Combustion in Aviation Gas Turbine Engines of University

Univers

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Cap

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The Influence of Fuel Properties on Threshold

Combustion in Aviation Gas Turbine Engines

Thesis presented for the degree

Doctor of Philosophy

in the Department Mechanical Engineering University of Cape Town

Author: Victor Burger Supervisor: Professor Andrew Yates

February 2017

The copyright of this thesis vests in the author. No quotation from it or information derived from it is to be published without full acknowledgement of the source. The thesis is to be used for private study or non-commercial research purposes only.

Published by the University of Cape Town (UCT) in terms of the non-exclusive license granted to UCT by the author.

Univers

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Abstract

This body of work investigated the influence of alternative jet fuel properties on aviation

gas turbine performance at threshold combustor operating conditions. It focused on

altitude blowout performance and was in part motivated by results that were

encountered during an aviation industry evaluation of synthetic kerosene that complied

with the Jet A-1 specification, but differed from the fuel that was used as a reference in

terms of some significant properties. As a consequence the relative impact of physical

properties and reaction chemistry properties were of primary interest in this study.

The thesis considered the potential to blend a range of different alternative jet fuel

formulations which exhibited independent variations in properties relating to

evaporation and reaction behaviour whilst still conforming to legislated physical fuel

specifications. It further explored the potential for said variations having a detectable

and significant influence on the simulated high altitude extinction behaviour in a

representative aviation gas turbine combustor. Based on the findings, appropriate

metrics were suggested for scientifically quantifying the appropriate properties and

conclusions were drawn about the potential impact of alternative jet fuel properties on

blowout performance.

These subjects were addressed primarily through the theoretical analyses of targeted

experimental programmes. The experimental design adopted a novel approach of

formulating eight test fuels to reflect real-world alternative fuel compositions while still

enabling a targeted evaluation of the influences of both physical and chemical reaction

properties. A detailed characterisation was performed of the test fuels’ physical and

reaction properties. The extinction and spray behaviours of the fuels were then

evaluated in a laboratory scale combustor featuring dual-swirl geometry and a single

prefilming airblast atomiser. The various experimental data sets were interpreted within

the context of a theoretical model analysis. In doing so the relative performance of

alternative jet fuel formulations under laboratory burner conditions were translated to

predict relative real world altitude performance. This approach was validated against

aforementioned industry evaluation results and demonstrated to be consistent.

A technically defensible explanation was provided for the previously unexplored

anomalous altitude extinction results that were observed during the industry evaluation

of synthetic jet fuel. A conclusive case was made for the extinction limit differences

having been caused by the relative differences in chemical ignition delays of the fuels.

ii

The probability of volatility (distillation profile) and fuel physical properties playing a

significant role in the impaired altitude performance was discredited. Evaporation-

controlled combustion efficiency was, however, shown to become a significant factor at

low air mass flow rates or when the fuel evaporation is compromised.

The influence of flame speed and chemical ignition delays were investigated. Laminar

flame speed was shown not to correlate with LBO, discrediting its use as a proxy for

reaction rate. The study showed a correlation between the lean blowout behaviour of

jet fuels and the ignition delays associated with their derived cetane numbers.

Additionally, there was substantive support indicating that an even stronger correlation

could be obtained by operating the IQT™ device that is used to measure these delays at

an elevated temperature.

The thesis makes a contribution towards the development of both technical

understanding and practical tools for evaluating the potential operating limits of

alternative jet fuel formulations.

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Acknowledgements

I wish to thank the following people:

Professor Andy Yates, for his indispensable guidance, encouragement and

support.

Dr Carl Viljoen for playing a key role in initiating the collaboration with German

Aerospace Centre (DLR) and for trying his best to educate a mechanical engineer

in the dark art of chemistry.

Dr Mariam Ajam and Dr Rina van der Westhuizen for their assistance with the

fuel matrix characterisation.

Sasol Energy Technology, headed by Paul Morgan, for the support and funding of

this project.

The DLR Institute of Combustion Technology, headed Prof. Dr Manfred Aigner,

and especially Dr Thomas Mosbach for hosting me in Stuttgart.

Dr Clemens Naumann for the shock tube ignition delay characterisation of the

test fuels.

Dr William O'Loughlin and Jasper Grohmann, for their assistance during the

model-combustor test measurements.

My parents for their love, support, and patiently explaining organic chemistry and

gas chromatography to me.

This thesis is dedicated to my wife, Zelda, for her love, support, and patience without

which I would not have been able to conclude this work.

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Contents

Abstract ................................................................................................................................. i

Acknowledgements ............................................................................................................ iii

List of Figures ...................................................................................................................... vi

List of Tables ........................................................................................................................ x

Acronyms ............................................................................................................................ xi

Nomenclature .................................................................................................................... xii

Subscripts ........................................................................................................................... xiv

1 Introduction ................................................................................................................. 1

1.1 Thesis structure ...................................................................................................... 2

1.2 FSJF certification: Altitude extinction findings ...................................................... 2

2 Preliminary Theoretical Assessment and Project Plan ................................................ 8

2.1 Combustion efficiency ........................................................................................... 8

2.1.1 Evaporation-controlled system ....................................................................... 9

2.1.2 Mixing-controlled system .............................................................................. 11

2.1.3 Reaction-controlled system .......................................................................... 12

2.2 Theoretical flight envelope .................................................................................. 14

2.2.1 Combustion efficiency influences ................................................................. 17

2.3 Hypotheses and project plan ............................................................................... 20

3 Literature Review ...................................................................................................... 23

4 Theoretical Assessment ............................................................................................. 34

4.1 Burning velocity model ........................................................................................ 34

4.2 Stirred Reactor model .......................................................................................... 36

4.3 Evaporation-controlled stability limits ................................................................ 41

4.4 Combustion Efficiencies ....................................................................................... 44

4.4.1 Evaporation-controlled efficiency ................................................................. 44

4.4.2 Mixing-controlled efficiency ......................................................................... 45

v

4.4.3 Reaction-controlled efficiency ...................................................................... 46

5 Experimental Design .................................................................................................. 49

5.1 Test fuel matrix .................................................................................................... 49

5.2 Physical characterisation of test fuel matrix ....................................................... 51

5.3 Chemical characterisation of test fuel matrix ..................................................... 51

5.4 Gas turbine model-combustor: LBO evaluation .................................................. 53

5.5 Gas turbine model-combustor: fuel spray evaluation ......................................... 56

6 Results ........................................................................................................................ 58

6.1 Physical characterisation results .......................................................................... 58

6.2 Chemical characterisation results ........................................................................ 58

6.3 Gas turbine model-combustor: LBO evaluation results ...................................... 64

6.4 Gas turbine model-combustor: fuel spray evaluation results ............................. 66

7 Analysis and Discussion of Results ............................................................................ 69

7.1 Spray and evaporation-controlled system analysis of LBO results ..................... 69

7.2 Reaction-controlled system analysis of LBO results ............................................ 75

7.3 Interpretation of the experimental results .......................................................... 81

8 Summary, Conclusions, and Recommendations ....................................................... 87

Appendix A: Evaporation-controlled System .................................................................. A.1

Appendix B: Comprehensive Fuel Property Analysis Results .......................................... B.1

Appendix C: Shock Tube Ignition Delay Results .............................................................. C.1

Appendix D: PDA Spray Measurement Results ...............................................................D.1

Appendix E: Modelled LBO Air Flow Proportionality Constants ..................................... E.1

References

vi

List of Figures

Figure 1.1: Mach number carpet - Relative performance of AVTUR and SFSAK ................. 4

Figure 1.2: ASTM D86 distillation curves for AVTUR and SFSAK test fuels .......................... 5

Figure 1.3: Lean blowout performance of JP-5 and FSJF relative to Jet A reference fuel ... 6

Figure 1.4: Distillation curves for Jet A, FSJF, and JP-5 test fuels ........................................ 7

Figure 2.1: Evaporation rate curves for kerosene and JP-4 ................................................. 9

Figure 2.2: Comparison of expressions for evaporation efficiency ................................... 11

Figure 2.3: The well-stirred reactor concept ..................................................................... 13

Figure 2.4: Standard atmosphere temperature, pressure and density relative to MSL values ............................................................................................................... 15

Figure 2.5: Typical idealised sub-sonic flight envelope...................................................... 16

Figure 2.6: Theoretical thrust limit, constant evaporation- and reaction-controlled combustion efficiency traces in the altitude-Mach number domain .............. 20

Figure 2.7: Schematic representation of thesis structure ................................................. 21

Figure 3.1: Ignition delay dataset correlation results with D/U and IDT as factors .......... 29

Figure 3.2: Derived cetane number against equivalence ratio at LBO .............................. 30

Figure 3.3: CPT index correlation for diffusion flame extinction ....................................... 31

Figure 3.4: Comparison of DCN to measured LTHR ........................................................... 32

Figure 4.1: Volumetric flow rate vs. efficiency in a typical well-stirred reactor ................ 38

Figure 4.2: Illustrative blowout curves ............................................................................... 39

Figure 4.3: Illustrative blowout curves for Jet A-1 and FSJF .............................................. 41

Figure 4.4: Illustrative reaction and evaporation limited extinction blowout curves ....... 43

Figure 4.5: Extinction performance for simulated wind milling operating condition, schematic representation of the result ........................................................... 43

Figure 4.6: Illustrative timescales in an evaporation limited system with fixed inlet pressure and temperature ............................................................................... 45

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Figure 4.7: Illustrative timescales in a mixing limited system with fixed inlet pressure and temperature ..................................................................................................... 46

Figure 4.8: Illustrative timescales in a chemical reaction limited system with fixed inlet pressure and temperature. .............................................................................. 47

Figure 5.1: Laboratory-scale model burner nozzle (configuration B) ................................ 54

Figure 5.2: Laboratory-scale model-combustor layout ..................................................... 55

Figure 5.3: Model-combustor setup .................................................................................. 56

Figure 5.4: Reference Jet A-1 flame ................................................................................... 56

Figure 6.1: Test fuel volatility: distillation profiles. ............................................................ 59

Figure 6.2: Test fuel CJF: Shock tube measured and calculated ignition delay results ..... 61

Figure 6.3: Modelled ignition delay results for the full tests fuel matrix .......................... 62

Figure 6.4: Comparison of IQT™ and shock tube ignition delays ...................................... 63

Figure 6.5: Model-combustor LBO results - burner configuration A, 323 K air pre-heat .. 65

Figure 6.6: Model-combustor LBO results - burner configuration A, 413 K air pre-heat .. 65

Figure 6.7: Model-combustor LBO results - burner configuration B, 323 K air pre-heat .. 66

Figure 6.8: Radial SMD profiles (ɸ = 0.6, 𝑚𝐴 = 6.5 g/s, z = 15 mm from exit plane) ......... 67

Figure 6.9: Radial SMD profiles (ɸ = 0.6, 𝑚𝐴 = 6.5 g/s, z = 35 mm from exit plane) ......... 68

Figure 6.10: Radial SMD profiles (ɸ = 0.6, 𝑚𝐴 = 2.2 g/s, z = 15 mm from exit plane) .. 68

Figure 7.1: Correlation of experimental and theoretical stoichiometric values (effect of proportional representation of configurations A and B) ................................. 74

Figure 7.2: Correlation of experimental and theoretical stoichiometric values (75% configuration B and 25% configuration A) ....................................................... 74

Figure 7.3: Proportionality constants for modelled combustor airflow based on LBO test data - burner configuration B, 323 K air pre-heat ........................................... 76

Figure 7.4: Modified proportionality constants (Kθ) for modelled combustor airflow based on LBO test data - burner configuration A, 323 K air pre-heat ............. 78

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Figure 7.5: Modified proportionality constants (Kθ) for modelled combustor airflow based on LBO test data - burner configuration B, 323 K air pre-heat ............. 78

Figure 7.6: Correlation of relative LBO proportionality constants (KAve) and intermediate temperature shock tube ignition delay (ID833 K) .............................................. 81

Figure 7.7: Theoretical relative reaction-controlled extinction limits ............................... 83

Figure 7.8: Theoretical relative evaporation-controlled extinction limits ......................... 84

Figure 7.9: Theoretical relative evaporation- and reaction-controlled extinction limits .. 86

Figure A.1: Evaporation rate curves for kerosene and JP-4 ............................................. A.2

Figure A.2: Drop size vs time ignoring heat-up ................................................................ A.2

Figure A.3: Droplet shell ................................................................................................... A.2

Figure A.4: Comparison of the direct and exponential expressions for evaporation efficiency ......................................................................................................... A.6

Figure C.1: Test fuel CJF: Shock tube ignition delay results .............................................. C.1

Figure C.2: Test fuel CJF/D: Shock tube ignition delay results .......................................... C.1

Figure C.3: Test fuel SJF1: Shock tube ignition delay results ............................................ C.2

Figure C.4: Test fuel SJF2: Shock tube ignition delay results ............................................ C.2

Figure C.5: Test fuel SJF2+: Shock tube ignition delay results .......................................... C.3

Figure C.6: Test fuel LTSK: Shock tube ignition delay results............................................ C.3

Figure C.7: Test fuel HTSK: Shock tube ignition delay results ........................................... C.4

Figure C.8: Test fuel HN: Shock tube ignition delay results .............................................. C.4

Figure D.1: Radial SMD profiles (ɸ = ɸLBO+5%, 𝑚𝐴 = 2.2 g/s, z = 15 mm from exit) ........D.3






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Figure D.7: Radial SMD profiles (ɸ = 0.6, 𝑚𝐴 = 2.2 g/s, z = 15 mm from exit) .................D.6



Figure D.10: Radial SMD profiles (ɸ = 0.6, 𝑚𝐴 = 6.5 g/s, z = 15 mm from exit) ...............D.7



Figure E1: Proportionality constants for modelled combustor airflow based on LBO test data - burner configuration A, 323 K air pre-heat .......................................... E.2

Figure E2: Proportionality constants for modelled combustor airflow based on LBO test data - burner configuration A, 413 K air pre-heat .......................................... E.2

Figure E3: Proportionality constants for modelled combustor airflow based on LBO test data - burner configuration B, 323 K air pre-heat .......................................... E.3

Figure E4: Modified proportionality constants (Kθ) for modelled combustor airflow based on LBO test data - burner configuration A, 323 K air pre-heat ............ E.3

Figure E5: Modified proportionality constants (Kθ) for modelled combustor airflow based on LBO test data - burner configuration A, 413 K air pre-heat ............ E.4

Figure E6: Modified proportionality constants (Kθ) for modelled combustor airflow based on LBO test data - burner configuration B, 323 K air pre-heat ............ E.4

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List of Tables

Table 2.1: International Standard Atmosphere, MSL Conditions ..................................... 15

Table 5.1: Test fuel matrix ................................................................................................ 50

Table 6.1: Test fuel characterisation: key physical properties ......................................... 58

Table 6.2: Test fuel characterisation: GCxGC speciation, DCN, LFS and shock tube ignition delays .................................................................................................. 59

Table 6.3: Correlation analysis of IQT™ and shock tube ignition delay results ................ 64

Table 7.1: PDA spray measurements: Relative SMD values averaged per axial measurement position ..................................................................................... 70

Table 7.2: Average relative effective evaporation constants, SMDs, and evaporation- controlled combustion efficiencies of the test fuel matrix. ............................ 71

Table 7.3: Correlation analysis of relative spray and evaporation metrics and the relative experimental PDA results. ................................................................................ 72

Table 7.4: Correlation analysis of normalised LBO proportionality constants (K) and modified proportionality constants (Kθ) .......................................................... 79

Table 7.5: Correlation analysis of average relative LBO behaviour, fuel spray, LFS and ignition delays .................................................................................................. 80

Table C.1: Arrhenius function coefficients for modelled ignition delay .......................... C.5

Table D.1: PDA spray measurements: Relative SMD values averaged per axial measurement position ....................................................................................D.2

Table E1: Normalised LBO proportionality constants and modified LBO proportionality constants ......................................................................................................... E.1

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Acronyms

AFT (p 38) adiabatic flame temperature

ALR (p 18) air/liquid mass ratio

AVTUR (p 3) aviation turbine fuel – reference Jet A-1

CJF (p 49) crude-derived Jet A-1

CJF/D (p 49) crude-derived Jet A-1 + 50% n-dodecane

DCN (p 27) derived cetane number

DEF STAN (p 2) UK defence standard

DLR (p 51) German Aerospace Centre

FAR (p 50) fuel/air mass ratio

FID (p 50) flame ionisation detector

FSJF (p 2) fully synthetic jet fuel

FT (p 48) Fischer-Tropsch

GC (p 50) gas chromatography

GCxGC (p 50) two-dimensional gas chromatography

GTL (p 26) gas to liquids

HCPP (p 49) hydrogenated cat-poly petrol

HTSK (p 49) synthetic paraffinic kerosene (SPK)

HN (p 49) heavy naphtha refinery stream

ICAO (p 14) International Civil Aviation Organisation

ISA (p 14) international standard atmosphere

IQT™ (p 30) ignition quality tester (manufactured by AET, Canada)

LBO (p 1) lean blowout

LFS (p 25) laminar flame speed

LTHR (p 30) low temperature heat release

LTSK (p 49) experimental GTL kerosene

MSL (p 14) mean sea level

NTC (p 59) negative temperature coefficient

OEM (p 2) original equipment manufacturer

PDA (p 50) phase Doppler anemometry

PIV (p 53) particle image velocimetry

SFSAK (p 3) Sasol fully synthetic aviation kerosene (see SJF1)

SJF1 (p 49) FSJF employed during certification process

SJF2 (p 49) commercial FSJF

SJF2+ (p 49) commercial FSJF + 1.5% HCPP

xii

SMD (p18) Sauter mean diameter

SPK (p 26) synthetic paraffinic kerosene

TOF-MS (p 50) time–of–flight mass spectrometry

Nomenclature

Af flame area, [m2]

Aref combustor reference area, [m2]

B constant (eq. 2.7)

d dynamic head, [m]

dyF dimensionless fuel disappearance fraction

dq quench distance, [m]

D instantaneous droplet diameter, characteristic diameter (eq. 4.6) [m]

Da Damköhler number

D0 starting droplet diameter, [m]

DP prefilmer diameter, [m]

Dref width of combustor casing, [m]

D32 Sauter mean diameter, [m]

fc fraction of airflow involved in combustion

𝐺𝑦 axial momentum flux

𝐺𝜃 angular momentum flux

h height above sea level, [km]

H lower specific energy of fuel, [J/kg]

𝑖 reaction order exponent

𝑗 reaction order exponent

K proportionality constant

Kθ modified proportionality constant

K’CJF reaction-controlled proportionality constant for reference fuel CJF

K”CJF evaporation-controlled proportionality constant for reference fuel CJF

𝑙 mixing length, [m]

𝑚 mass of fuel droplet, [kg]

𝑚0 starting droplet mass, [kg]

�̇� rate of fuel evaporation / mass change rate, [kg/s]

�̇�𝑓 rate of fuel evaporation per unit surface area, [kg/m2s]

n number of droplets

�̇� molar air flow rate, [moles/s]

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P pressure, [Pa]

ΔPL combustor liner pressure differential, [Pa]

q fuel/air mass ratio

qref reference dynamic pressure, [Pa]

R gas constant / characteristic radial distance (eq. 5.1)

Re drop Reynolds number

Ru universal gas constant

Sa annular geometric swirl number

Sc centre geometric swirl number

SL laminar flame speed, [m/s]

St small scale turbulent flame speed, [m/s]

ST large scale turbulent flame speed, [m/s]

t time, [s]

te droplet lifetime, [s]

Δtr residence time of air in combustor, [s]

T temperature, [K]

ΔT temperature rise due to combustion, [K]

�̅� rms component of fluctuating velocity, [m/s]

U velocity, [m/s]

Uj turbulent jet velocity, [m/s]

Uref combustor reference velocity, [m/s]

V volume, [m3]

v velocity, [m/s]

𝑥𝐹 fuel mole fraction

𝑥𝑂 oxygen mole fraction

𝑥𝑆 stoichiometric fuel mole fraction

YF overall fuel fraction consumed

α thermal diffusivity

𝜖 turbulent exchange coefficient

ɸ equivalence ratio

λ evaporation constant, [m2/s]

λeff effective evaporation constant, [m2/s]

𝜈 kinematic viscosity, [m2/s]

𝜇 dynamic viscosity, [kg/m.s]

ηc combustion efficiency

xiv

ηe combustion efficiency (evaporation-controlled)

ηm combustion efficiency (mixing-controlled)

ηθ combustion efficiency (reaction-controlled)

τ delay

𝜏𝑐 constant delay offset

τe evaporation timescale

τm mixing timescale

τres residence timescale

τθ reaction timescale

𝜌 density, [kg/m3]

𝜎 surface tension, [kg/s2]

Subscripts

ave average value

A air

c combustion zone value

F fuel

L liquid

ov overall value

(rel) relative to reference value

sl sea level

0 initial value

3 combustor inlet value

1

1 Introduction

Gas turbines dominate aviation propulsion and while they have traditionally always

been considered to be omnivorous fuel consumers, the demands of modern aviation

has resulted in a much more constrained fuel tolerance. Aviation gas turbines need to

operate over a wide range of ambient conditions while attaining ever increasing

efficiency and emission targets. In order to satisfy these demands jet fuel is subjected to

performance-based specifications which have traditionally allowed any combination of

hydrocarbons that satisfied the required performance, provided that they were

petroleum-derived. Significant performance properties include energy content, smoke

point, thermal stability, lubricity, volatility, non-corrosivity and cleanliness. Since the

1940s the specifications have focused predominantly on distillation, volatility, freeze

point, thermal stability and volume yield [1, 2]. In addition, handling and safety

requirements constrain a number of fuel properties. The consequence is that modern-

day aviation fuel is controlled by very explicit specifications that evolved out of a global

industry consensus approach. Environmental and energy security considerations have

stimulated interest in alternatives to petroleum-derived jet fuel which in turn resulted in

an industry-wide review of the complete list of jet-fuel specifications. Sasol played a key

role in this process as the world’s first producer of approved and certified semi-synthetic

and fully-synthetic jet fuels [3].

It is noteworthy that no direct or indirect fuel autoignition specifications are applied to

jet fuel, presumably due to the argument that under normal combustion conditions

chemical reaction timescales are negligible relative to spray vaporization time scales [4].

The combustion performance of practical aircraft combustors is known to be influenced

more by physical processes of heat transfer, mass transfer, thermodynamics, gas

dynamics and fluid dynamics than by chemical processes. While chemical reaction rates

are acknowledged as being of great importance, they are largely disregarded in high

temperature flames due to being considered to be relatively rapid and not rate

controlling. Emphasis is placed on large scale mixing and the interdiffusion of fuel and

air as the rate controlling steps. Chemical processes are, however, known to be of

particular importance for the formation of pollutant emissions and for the lean light-off

and lean blowout (LBO or weak extinction) limits at high flight altitudes [5].

Literature acknowledges the importance of chemical reaction timescales in the

combustion of well atomised fuels at low pressures and lean equivalence ratios as well

as under threshold combustion conditions such as encountered during altitude blowout

Introduction

2

and relight. Increased interest in alternative aviation fuel formulations, and the related

opportunity to influence chemical reaction rates to a greater degree than traditionally

possible in the case of petroleum-derived jet fuel, has prompted examination of the

influence of physical and chemical fuel properties on gas turbine combustion behaviour.

A number of studies have reported ignition and extinction behaviour being influenced

by fuel formulation to varying degrees although it is difficult to separate the influence of

fuel formulation on chemical reaction and fuel evaporation and mixing. These studies

employed experimental setups ranging from laboratory scale combustors to full scale

gas turbines and test fuels ranging from single component fuels to full boiling range jet

fuel formulations. Findings ranged from reports of appreciable differences to relative

insensitivity to fuel formulation [6, 7, 8, 9, 10, 11].

The potential impact of fuel formulation on threshold combustion was also highlighted

by altitude blowout results that were encountered during the evaluation of synthetic

kerosene (Sasol fully synthetic jet fuel or FSJF) as a Jet A-1 aviation turbine fuel under

DEF STAN 91-91 [12, 13].

1.1 Thesis structure

The thesis commences with a summary of the altitude blowout experience that

emerged during the FSJF certification process. This is followed by a chapter entailing a

preliminary theoretical assessment of combustion efficiency and its impact on the flight

envelope. The context of the FSJF results and the theoretical assessment provide the

foundation for the formulation of three thesis hypotheses and the project plan to

address these. A review of relevant literature follows in order to provide context and

background. Based on the guidance provided by the preliminary theory and literature, a

more detailed and focused theoretical assessment is presented. The experimental

design that was formulated to address the hypotheses is presented next. The

experimental results follow as does a subsequent analysis and discussion chapter aimed

at interpreting the results in terms of the hypotheses. Finally some concluding remarks

and recommendations are presented.

1.2 FSJF certification: Altitude extinction findings

During the FSJF certification process, Southwest Research Institute (SwRI) assessed the

results from detailed laboratory testing and submitted the findings to the engine

original equipment manufacturers (OEMs), the British MOD Aviation Fuels Committee,

the Aviation Fuels Group of the Coordinating Research Council, and the ASTM J1

Chapter 1

3

Aviation Fuels Subcommittee. Following the evaluation of the fuel properties and

characteristics, as detailed in a comprehensive report by Moses and Wilson [12], the

OEMs requested engine and combustor tests to confirm no adverse effects on engine

performance and operation. This resulted in a series of test programs being conducted

and or monitored by the OEMs focusing on engine performance, endurance, emissions,

atomisation, ignition and relight, and lean blowout. Moses compiled a report detailing

the results and interpretation of the various engine test programmes, which was

instrumental in the eventual approval of FSJF under DEF STAN 91-91 [13]. The findings

from these test reports contained the following aspects which related specifically to

altitude ignition and extinction.

Rolls-Royce plc. was entrusted with conducting a combustor rig-based evaluation of

ground ambient cold day relight and extinction performance. The test program was

executed at Smiths Ltd’s high altitude facility in Burnley, UK. A full annular combustion

test system representative of a modern aero gas turbine was used in a configuration

equipped with airblast fuel spray nozzles. A reference test fuel (AVTUR) was evaluated

alongside the FSJF, which was referred to in the report as “Sasol Fully Synthetic Aviation

Kerosene” or SFSAK. Ignition, stability (extinction limit) and blowout tests were

conducted at several altitude pressures to map the likely operating envelope. A detailed

summary of the test methodology is contained in the full test report [14]. Both the

reference AVTUR and SFSAK were mapped over a range of operating conditions with

ignition, extinction, and blowout loops being compared at the same conditions (e.g.

same time to ignition and number of sectors lit/extinguished). The ignition and

extinction performance was translated to an altitude-Mach number realm by applying

typical windmilling performance to calculate altitude and Mach numbers for each

combustor pressure and air mass flow combination. Results from more than 200 test

points are presented in Figure 1.1 (from the test report) which includes three data sets:

one ground ambient cold day case and two altitude cases. Although the ground ambient

cold day results are not a windmilling relight condition, it was included in order to

present all the data on a single graph.

The figure reflects the blowout extinction and ignition conditions for SFSAK and AVTUR.

Some differences between the fuels were noted in the high altitude, high sub-sonic

Mach number region. The most distinct differences were recorded in the relative

blowout extinction performance of the two fuels. The lean and rich extinction

boundaries of the two fuels were close, but differences increased towards the blowout

point with an almost 6% difference in relative air flow at first sector blowout and more

Introduction

4

than 10% difference in final sector blowout. The SFSAK blowout points occurred at

lower combustor mass flows than those recorded by the AVTUR fuel. This translated

into the significant difference in blowout extinction points of the two fuels in the higher

altitude and Mach number regime of the windmilling envelope shown in Figure 1.1.

Only marginal ignition boundary differences were recorded and were considered to be

within the test method accuracy of 0.005 Fuel-to-Air ratio. The SFSAK toe ignition point

occurred at a 2.5% lower air mass flow than for AVTUR.

Point with lower air mass flow than Toe point; Ignition Toe point for AVTUR;

Ignition Toe point for SFSAK; Blowout point for AVTUR; Blowout point for SFSAK

Figure 1.1: Mach number carpet - Relative performance of AVTUR and SFSAK (from original Rolls-Royce report [14])

The report focused on differences in boiling point distribution between the two fuels

(Figure 1.2) and identified it as a potential cause for the blowout differences. It was,

however, acknowledged that “the possibility of a chemical issue could not be ruled out.”

In light of these results the OEMs recommended further lean blowout tests at lower

Chapter 1

5

altitude conditions that are more relevant to descent where throttle manoeuvres are

expected to be more likely to cause lean combustion.

Figure 1.2: ASTM D86 distillation curves for AVTUR and SFSAK test fuels (from original Rolls-Royce report [14])

In response to the issue raised by Rolls-Royce, Honeywell Aerospace conducted low-

temperature atomisation (at -40°C) and lean blowout tests. The atomisation of the FSJF

was reported to be as good as or better than the Jet A reference fuel for both airblast

and pressure type fuel atomisers. This was ascribed to the lower specific gravity and

viscosity of FSJF and it was concluded that FSJF would provide well-defined spray

distribution and droplet size for reliable cold starting [15]

Lean stability or lean blowout testing was conducted at representative flight and ground

deceleration conditions from sea level up to an altitude of 25000 ft. A full scale reverse-

flow annular combustor rig with dual-orifice fuel atomisers was employed. Three fuels

were evaluated, the FSJF test fuel, Jet A (ASTM D 1655), and weathered JP-5 (MIL-DTL-

5624) which had some volatile components removed and therefore exhibited a flat

distillation profile similar to the FSJF . Blowout points were obtained by reducing fuel

flow after stable combustion had been established. Further detail of the test facility and

measurement procedure is contained in the report.

The results as shown in Figure 1.3 were reported “relative” to those obtained by the

Jet A reference and the authors of the report concluded that lean stability at sea level

was very similar for the three fuels. A “limited fuel effect” was noted at the higher

Introduction

6

simulated altitude (25000 ft.) test point were the FSJF results fell between those

recorded by JP-5 and the Jet A reference. The relative differences were considered to be

small and ordered according to the front-end boiling point distribution of the three

fuels. The distillation curves of the three fuels are presented in Figure 1.4. The JP-5

exhibited a similarly flat distillation profile to that of the FSJF, but while it had the same

flash point the overall distillation was shifted to higher boiling points and the front-end

volatility ranking is apparent. Honeywell concluded that the blowout performance

differences that were reported by Rolls-Royce were not due to chemistry and that FSJF

recorded no adverse effect on combustor lean stability characteristics. The differences

at higher altitude conditions were ascribed to boiling point distribution; volatility and

flash point in particular.

Figure 1.3: Lean blowout performance of JP-5 and FSJF relative to Jet A reference fuel (from original Honeywell Aerospace report [15])

Chapter 1

7

Figure 1.4: Distillation curves for Jet A, FSJF, and JP-5 test fuels (from original Honeywell Aerospace report [15])

While the two studies by Honeywell and Rolls-Royce provided the industry with

sufficient confidence to certify FSJF (with the specified minimum distillation gradient),

the results suggested that the relative influences of physical and chemical properties on

altitude ignition and extinction had not been addressed fully. In order to explore this

subject further it is necessary to first consider the theoretical constraints that apply to

combustion efficiency over the altitude-velocity operating envelope.

8

2 Preliminary Theoretical Assessment and Project Plan

In order to formulate a project hypothesis, it was necessary to subject the results from

the Rolls-Royce and Honeywell Aerospace evaluations to a preliminary theoretical root-

cause analysis. Aviation safety authorities such as the Federal Aviation Administration

and the European Aviation Safety Agency require the submission and validation of so-

called flight envelopes that define the ignition and extinction boundaries of combustion

in an aviation gas turbine. These flight envelopes are essentially the result of the

interplay between atmospheric conditions, hardware design and operation, and

combustion efficiency. In the following section simplified theoretical treatments of

combustion efficiency and flight envelopes are employed to examine the potential for

different fuel properties to influence flight envelope boundaries.

2.1 Combustion efficiency

Gas turbine combustors need to burn stably over a wide range of operating conditions

while maintaining overall combustion efficiencies in excess of 99 percent in order to

meet emissions regulations. Aviation gas turbines furthermore need to ignite readily

after an inflight flameout and attain primary zone combustion efficiencies of 75 to 80

percent as they accelerate back up to normal operation. Lower combustion efficiency

levels could result in rich extinction due to over fuelling in order to overcome the

narrow stability limits at altitude while higher levels could result in an unnecessarily

large combustor [16]. A well-established model for relating combustion efficiency to

combustor operating conditions and dimensions is based on the total time required for

combustion being equal to the sum of the times required for fuel evaporation, mixing of

the reactants (fuel vapour, air and combustion products) and the chemical reaction. The

time available for combustion is inversely proportional to the airflow rate, which means

that the combustion efficiency can be expressed as follows.

𝜂𝑐 = 𝑓 ((𝑎𝑖𝑟𝑓𝑙𝑜𝑤 𝑟𝑎𝑡𝑒)−1 (

1

𝑒𝑣𝑎𝑝𝑜𝑟𝑎𝑡𝑖𝑜𝑛 𝑟𝑎𝑡𝑒+

1

𝑚𝑖𝑥𝑖𝑛𝑔 𝑟𝑎𝑡𝑒+

1

𝑐ℎ𝑒𝑚𝑖𝑐𝑎𝑙 𝑟𝑒𝑎𝑐𝑡𝑖𝑜𝑛 𝑟𝑎𝑡𝑒)−1

) [2.1]

It is understood that under normal conditions the maximum heat release rate is

governed by either evaporation, mixing or chemical reaction rates, but under transient

conditions between either of these regimes the overall combustion efficiency could be

influenced by two of the three concurrently.

Chapter 2

9

2.1.1 Evaporation-controlled system

In a system where the mixing and reaction rates are sufficiently fast for the fuel

evaporation to be the rate-controlling step, the fuel is assumed to instantly mix and

burn with the surrounding air as soon as it evaporates. Under such conditions the

evaluation of evaporation and droplet lifetime is of interest since it determines the

residence time required for combustion to occur. An expanded treatment including all

relevant derivation is included in Appendix A.

The evaporation of a spherical droplet is accepted to consist of an unsteady state or

heat-up period during which the droplet temperature increases with time until the drop

attains its wet-bulb temperature. This transient period is followed by relatively steady-

state evaporation during which the droplet diameter, D, changes according to the so-

called “D2 law of droplet evaporation” that was first formulated by Godsave [17]. Typical

assumptions include: that the droplet is spherical, that the fuel is a pure single boiling

point liquid and that radiant heat transfer is negligible. The mathematical expression of

the law is provided in Equation 2.2 where λ represents the evaporation constant, t the

time elapsed and D and D0 the instantaneous and starting droplet diameters

respectively [18].

𝐷𝑜2 − 𝐷2 = 𝜆𝑡 [2.2]

Figure 2.1 shows this relationship between D2 and time as generated by Wood et al. for

kerosene and JP-4 droplets [19].The fuel-dependent evaporation constant corresponds

to the slope of the linear portion of the graph. During the initial heat-up stage the slope

of the graph is very small while the majority of the heat raises the droplet temperature.

Figure 2.1: Evaporation rate curves for kerosene and JP-4 (from Wood et al. [19])

Preliminary Theoretical Assessment and Project Plan

10

From Equation 2.2 it follows that the evaporation constant, or slope of the curves in

Figure 2.1, are expressed by

𝜆 = 𝑑

𝑑𝑡 (𝐷2) [2.3]

As detailed in Appendix A the average fuel spray evaporation rate �̇�𝑎𝑣𝑒, can be

expressed in terms of the air density 𝜌𝐴, an effective evaporation constant defined by

Chin and Lefebvre [20] λeff, air volume V, fuel/air ratio q, and the initial diameter, D0.

�̇�𝑎𝑣𝑒 = 𝜌𝐴𝜆𝑒𝑓𝑓𝑉𝑞

𝐷02⁄ [2.4]

This expression for average evaporation rate can be employed in determining the

combustion efficiency of a system in which the mixing and reaction rates are sufficiently

fast for the fuel evaporation to be the rate-controlling step. The combustion efficiency is

therefore governed by the ratio of the rate of fuel evaporation within the combustion

zone to the rate of fuel supply.

𝜂𝑒 = �̇�𝑎𝑣𝑒

𝑓𝑐𝑞𝑐�̇�𝐴⁄ [2.5]

The fraction of the total combustor airflow that takes part in combustion is represented

by fc and the combustion zone fuel/air ratio by qc. Substituting 2.4 into 2.5 yields:

𝜂𝑒 = 𝜌𝐴𝜆𝑒𝑓𝑓𝑉

𝐷02𝑓𝑐�̇�𝐴

⁄ [2.6]

Equation 2.6 expresses combustion efficiency in terms of the combustor operating

conditions (𝜌𝐴 , 𝜆𝑒𝑓𝑓 , �̇�𝐴), physical combustor dimensions (V), fuel spray atomiser

properties (D0) and fuel properties (D0 , 𝜆𝑒𝑓𝑓). The equation, however, allows the

evaporation efficiency to assume a value greater than 100%. This occurs when the time

required for complete evaporation is less than the available time and the fuel within the

primary recirculation zone is thus fully vaporised. Under these conditions the

combustion efficiency should be assigned a value of 100%. Lefebvre proposes the use of

a modified form similar to 2.7 in order to avoid the abrupt discontinuity [21]. The two

expressions are compared on the basis of combustion air residence in Figure 2.2. The

value of the shape-function constant, B, is typically chosen so that the efficiency derived

by 2.6 and 2.7 are coincident at the blowout efficiency of around 80%. It should be

noted that 2.7 can over or under predict the base expression 2.6 by up to 15%.

Chapter 2

11

𝜂𝑒 = 1 − exp (−𝐵𝜌𝐴𝜆𝑒𝑓𝑓𝑉


⁄ ) [2.7]

Figure 2.2: Comparison of the direct (eq. 2.6) and exponential (eq. 2.7) expressions for

evaporation efficiency

2.1.2 Mixing-controlled system

In his treatment of a system where the fuel evaporation and the chemical reaction

kinetics are assumed to be infinitely fast, Lefebvre expressed the combustion efficiency

as a function of the mixing and airflow rate [22].

𝜂𝑚 = 𝑓(𝑚𝑖𝑥𝑖𝑛𝑔 𝑟𝑎𝑡𝑒 𝑎𝑖𝑟 𝑓𝑙𝑜𝑤 𝑟𝑎𝑡𝑒⁄ ) [2.8]

The mixing rate of a turbulent air jet with the gas surrounding it is the product of the

eddy diffusivity, the mixing area, and the density gradient between the jet and the

surrounding gas. By assuming the eddy diffusivity to be proportional to the product of

the mixing length, 𝑙, and the turbulent jet velocity, 𝑈𝑗, the mixing rate can be expressed

as follows:


12

𝑚𝑖𝑥𝑖𝑛𝑔 𝑟𝑎𝑡𝑒 = (𝑒𝑑𝑑𝑦 𝑑𝑖𝑓𝑓𝑢𝑠𝑖𝑣𝑖𝑡𝑦)(𝑚𝑖𝑥𝑖𝑛𝑔 𝑎𝑟𝑒𝑎)(𝑑𝑒𝑛𝑠𝑖𝑡𝑦 𝑔𝑟𝑎𝑑𝑖𝑒𝑛𝑡)

= (𝑙𝑈𝑗)(𝑙2)(𝜌/𝑙)

= 𝜌𝑈𝑗𝑙2 [2.9]

In a combustor where the turbulent jet velocity is a function of the liner pressure

differential, ∆𝑃𝐿, and density: 𝑈𝑗 ∝ (∆𝑃𝐿 𝜌3⁄ )0.5, the mixing rate can be expressed as:

𝑚𝑖𝑥𝑖𝑛𝑔 𝑟𝑎𝑡𝑒 ∝ (𝜌3∆𝑃𝐿)0.5𝑙2

∝ (𝑃3𝑙2 𝑇3

0.5⁄ )(∆𝑃𝐿 𝑃3⁄ )0.5 [2.10]

Substituting the mixing rate expression into 2.8 and assuming the mixing length to be

proportional to the combustor size, 𝐴𝑟𝑒𝑓, the combustion efficiency in a mixing-

controlled system can be expressed as [22]:

𝜂𝑚 = 𝑓 ((𝑃3𝐴𝑟𝑒𝑓 �̇�𝐴𝑇30.5⁄ )(∆𝑃𝐿 𝑃3⁄ )0.5) [2.11]

2.1.3 Reaction-controlled system

There are several theoretical approaches which attempt to characterise combustion

efficiency in a system where chemical kinetics are considered to be rate-controlling and

evaporation and mixing rates are assumed to both be infinitely fast [23, 24, 25]. These

will be examined in greater detail later, but for this preliminary assessment the

commonly used well-stirred reactor provides an adequate representation.

The well-stirred reactor approximates the combustion zone as a perfectly stirred reactor

into which air and fuel flows at a constant rate and the reactants are mixed

instantaneously with all material in the reactor. Combustion products are assumed to

leave the reactor at a constant rate with temperature and compositional properties

identical to that of the reactor or zone. In Lefebvre’s treatment the governing reaction

rate equation (2.12) is presented without explanation [26] as is the case in his

referenced source of this equation, Longwell et al. [27].

𝜂𝑐𝜙�̇�𝐴 = 𝐶𝑐𝑓𝑉𝑇0.5 𝑒(−𝐸 𝑅𝑇⁄ )𝜌𝑛𝑥𝐹

𝑚𝑥𝑂𝑛−𝑚 2.12]

Chapter 2

13

Figure 2.3: The well-stirred reactor concept

A full derivation has, however, been presented by Strehlow [28] based on the

theoretical reactor volume depicted in Figure 2.3. The fuel supply in moles per second is

given by 𝐹𝑢𝑒𝑙 𝑆𝑢𝑝𝑝𝑙𝑦 = 𝜙𝑥𝑠�̇�, where 𝜙 is the molar equivalence ratio, 𝑥𝑠 is the

stoichiometric fuel mole fraction and �̇� is the air flow rate in moles per second.

Assuming a steady flow of premixed reactants into an unstirred reactor, the fuel in

volume element dV is consumed at a rate - 𝑑𝑑𝑡[𝐹]𝑑𝑉 where the fuel concentration, [𝐹], is

the number of fuel moles per unit volume. A dimensionless fuel disappearance fraction,

𝑑𝑦𝐹, can be defined as follows.

𝑑𝑦𝐹 = 𝐿𝑜𝑐𝑎𝑙 𝑓𝑢𝑒𝑙 𝑐𝑜𝑛𝑠𝑢𝑚𝑝𝑡𝑖𝑜𝑛 𝑟𝑎𝑡𝑒

𝑂𝑣𝑒𝑟𝑎𝑙𝑙 𝑓𝑢𝑒𝑙 𝑠𝑢𝑝𝑝𝑙𝑦 𝑟𝑎𝑡𝑒=

−𝑑

𝑑𝑡[𝐹]𝑑𝑉

𝜙𝑥𝑠�̇�

∴ 𝑑

𝑑𝑡[𝐹]𝑑𝑉 = −𝜙𝑥𝑠�̇�𝑑𝑦𝐹 [2.13]

The differential equation can be solved by grouping the variables and integrating over

the reactor volume as well as the fuel consumption profile.

∫𝑑𝑉

𝜙𝑥𝑠�̇�𝑉= −∫

𝑑𝑦𝐹𝑑

𝑑𝑡[𝐹]

𝑌𝐹

0 [2.14]

where the limit, 𝑌𝐹, is the overall fuel fraction consumed. The integral on the right is

difficult to evaluate due to the reaction rate, 𝑑

𝑑𝑡[𝐹], being a function of both

concentration and time. By assuming the reactor to be well-stirred containing a

homogeneous mixture of reactants and products, 𝑑

𝑑𝑡[𝐹] is assumed to be constant

which means that the integration yields:

𝑉

𝜙𝑥𝑠�̇�= −

𝑌𝐹𝑑

𝑑𝑡[𝐹]

[2.15]

∴ 𝑥𝑠�̇�

𝑉= −

1

𝜙𝑌𝐹

𝑑

𝑑𝑡[𝐹] [2.16]

mixture products

dV


14

The reaction rate can be expressed in terms of the fuel and oxidiser concentrations and

an Arrhenius temperature dependence.

Note: Strehlow uses n and m as reaction order coefficients, while Lefebvre uses m and

n-m notation. The fuel and oxidiser reaction order coefficients 𝑖 and 𝑗 are therefore

used here to avoid confusion.

𝑑

𝑑𝑡[𝐹] ∝ −[𝐹]𝑖 [𝑂]𝑗𝑒−𝐸 𝑅𝑇⁄ [2.17]

The initial concentrations of fuel and oxidiser in a constant-volume adiabatic system

may be written as [𝐹] = 𝑥𝐹 𝑃𝑜𝑅𝑢𝑇𝑜

and [𝑂] = 𝑥𝑂 𝑃𝑜𝑅𝑢𝑇𝑜

.

With �̇� ∝ �̇�𝐴 and 𝑌𝐹 ≡ 𝜂𝑐 = 𝜂𝜃, Equation 2.17 into 2.16 results in:

𝜂𝜃𝜙�̇�𝐴 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌𝑖+𝑗𝑥𝐹

𝑖 𝑥𝑂𝑗

[2.18]

Strehlow omitted a pre-exponential temperature dependence from the Arrhenius

expression which Laidler and Lefebvre included, resulting in the T0.5 discrepancy

between equations 2.12 and 2.18. This is revisited in Chapter 4.

2.2 Theoretical flight envelope

The operating flight envelope of an aviation gas turbine is shaped by atmospheric

conditions, hardware considerations and combustion efficiency limits. Subsonic civil

aircraft operate in the troposphere and lower stratosphere up to typical cruise altitudes

of approximately 38000 ft. (11.5 km) [29]. Knowledge of the vertical distribution of

temperature, pressure, density, viscosity and the speed of sound within the atmosphere

is essential for the modelling of flight conditions. The atmosphere, however, does not

remain constant at any particular place or time which necessitated the development of

a hypothetical approximation known as the standard atmosphere. The International

Organisation of Standardisation (ISO) publishes the International Standard Atmosphere,

ISA, while a number of authorities such as the International Civil Aviation Organisation

(ICAO) and the United States Government publish extensions and subsets of the same

atmospheric model. The models assume the air to be at rest (zero turbulence and no

wind) and to be free of any moisture or dust. Naturally significant differences are

possible between actual atmospheric conditions and the standard atmosphere

properties, which are based on average values recorded at a latitude of 45° N. Table 2.1

provides a summary of the mean sea level (MSL) values employed by the ISA.

Chapter 2

15

Table 2.1: International Standard Atmosphere, MSL Conditions [30]

Temperature Tsl 288.15 K Pressure Psl 101325 N.m-2

Density ρsl 1.225 kg.m-3

Speed of Sound asl 340.294 m.s-1

Gravity gsl 9.80665 m.s-2

Temperature, pressure and density decrease with altitude as shown in Figure 2.4. The

ISA assumes that temperature falls linearly at a rate of 6.5 K/km in the troposphere

below the tropopause (11 km above sea-level). Above 11 km the ISA temperature

remains constant at 216.65 K up to an altitude of 20 km. The pressure, altitude

relationship is given by Equation 2.19 where h is the height above sea-level in km, M is

the molar mass of air and Ru is the universal gas constant. From the ideal gas law density

is given by Equation 2.20.

𝑃 = 𝑃𝑠𝑙 (1 −6.5ℎ

𝑇𝑠𝑙⁄ )

9.81𝑀 6.5𝑅𝑢⁄

[2.19]

𝜌 = 𝑃 𝑅𝑇⁄ [2.20]

Figure 2.4: Standard atmosphere temperature, pressure and density relative to MSL

values


16

A typical sub-sonic flight altitude-Mach number envelope, also sometimes referred to as

a “doghouse plot”, is shown in Figure 2.5. Boundary AB is governed by the minimum

aircraft speed based on either engine thrust or the minimum stall speed. BC is

determined by the operating ceiling of the aircraft and CD is governed by maximum

flight Mach number while DE is determined by the maximum engine thrust. The BCDE

boundary is of particular interest as far as the differences in altitude blowout and relight

behaviour by different fuels are concerned. Although the thermodynamic conditions in

the combustor are strongly influenced by the atmospheric temperature and pressure,

flight speed also plays a significant role as do design parameters such as bypass ratio,

power offtake and exhaust configuration. A simplified approach has been followed to

derive combustor temperatures, pressure and mass flow rates corresponding to the

flight operating envelope.

Figure 2.5: Typical idealised sub-sonic flight envelope

Since the greatest disparity between test fuels in the Rolls-Royce evaluation occurred

under blowout extinction conditions, this condition was simulated in this preliminary

assessment. Conditions prior to in-flight flame-out, or extinction were approximated as

follows. Positions 0, 2, and 3 refer to ambient, compressor inlet, and compressor outlet

positions respectively. For a given compressor ratio, 𝑟 = 𝑃3 𝑃0⁄ , the isentropic

temperature ratio from the compressor inlet to combustor inlet is given by:

Chapter 2

17

𝑇3is

𝑇2= 𝑟(𝛾−1) 𝛾⁄ [2.21]

Isentropic compressor efficiencies in modern aircraft gas turbines are typically around

90% [31] and treating the fluid as an ideal gas it can be expressed as

𝜂𝑐𝑜𝑚𝑝 = 𝑇3is− 𝑇2

𝑇3− 𝑇2 [2.22]

The actual temperature ratio over the compressor is thus given by

𝑇3

𝑇2= 1 +

𝑟(𝛾−1) 𝛾⁄ − 1

𝜂𝑐𝑜𝑚𝑝 [2.23]

The stagnation pressure can be calculated via the isentropic relationship

𝑃0 = 𝑃 (1 +𝛾−1

2𝑀2)

𝛾 (𝛾−1)⁄

[2.24]

2.2.1 Combustion efficiency influences

Boundary BCDE in Figure 2.5 can be impacted by combustion efficiency and therefore

fuel influences on combustion efficiency can potentially also delimit the envelope

boundary. This potential impact was investigated by applying the modelled evaporation-

mixing- and reaction-controlled efficiency expressions to the idealised altitude, Mach

number model. Physical hardware parameters and proportionality constants were

arbitrarily chosen to illustrate potential envelope limits. The mixing-controlled efficiency

as defined in Equation 2.11 is insensitive to flight conditions and does not impact

boundary BCDE. This is revisited in Chapter 4.

Equation 2.6 was employed to model the evaporation-controlled efficiency over the

operating altitude temperature, pressure, and air mass flow range.



⁄ [2.6]

Combustion efficiencies of at least 75 to 80% are required to sustain combustion and

prevent flameout [16]. A combustion efficiency limit of 75% was therefore used to

calculate the corresponding maximum flight Mach numbers. Chin and Lefebvre [20]

presented plots of effective evaporation constants versus boiling point for various

values of droplet size and velocity at three pressures and three ambient temperatures.

They acknowledged that while a single fuel property cannot fully describe evaporation

characteristics of any given fuel, average boiling point has the benefit of being directly


18

related to vapour pressure and fuel volatility. Using these three plots, Equation 2.25 was

derived empirically for calculating the effective evaporation constant. The omission of

droplet size and velocity was deemed to be an acceptable simplification for an initial

investigation. The fuel boiling point was treated as constant for a single fuel model and

is revisited in Section 7.3 for the evaluation of multiple fuels with variable boiling point

temperatures.

𝜆𝑒𝑓𝑓 ∝ 𝑃0.09 𝑇1.8 [2.25]

The majority gas turbines employ prefilming airblast atomisers combined with strong

swirling flow to ensure fine atomisation and dispersion of droplets throughout the

combustion zone. Rizkala and Lefebvre [32] concluded that for fluids such as kerosene

with relatively low viscosities the mean droplet size is primarily influenced by surface

tension, air velocity and air density while viscosity plays and independent role. They

proposed a two-term expression for Sauter mean diameter, SMD, (Equation 2.26) based

on analysis of experimental data where the first term is dominated by surface tension

and air momentum and the second by liquid viscosity. Lefebvre subsequently presented

a modified version (Equation 2.27) with experimentally determined burner-specific

constants A and B [33].

𝑆𝑀𝐷 = 3.33 × 10−3(𝜎𝜌𝐿𝐷𝑃)

0.5

𝜌𝐴𝑈𝐴(1 +

1

𝐴𝐿𝑅) + 13.0 × 10−3 (

𝜇𝐿2

𝜎𝜌𝐿)0.425

𝐷𝑃0.575 (1 +

1

𝐴𝐿𝑅)2

[2.26]

𝑆𝑀𝐷 = 𝐴 (𝜎

𝜌𝐴𝑈𝐴2𝐷𝑃)0.5

(1 +1

𝐴𝐿𝑅) + 𝐵 (

𝜇𝐿2

𝜎𝜌𝐿𝐷𝑃)0.5

(1 +1

𝐴𝐿𝑅)2

[2.27]

DP is the prefilmer diameter, 𝑈𝐴 the air velocity, and ALR the air/liquid mass ratio. For

low viscosity liquids such as water and kerosene the first term predominates resulting

in 2.28.

𝑆𝑀𝐷 = 𝐷32 ∝ (𝜎


(1 +1

𝐴𝐿𝑅) [2.28]

With 1 +1

𝐴𝐿𝑅≈ 1 +

𝜙

14.6≈ 1 and assuming constant hardware dimensions as well as fuel

properties, Equation 2.6 simplifies to:

𝜂𝑒 ∝ 𝑃1.09𝑇0.8 / �̇�𝐴 [2.29]

Applying the atmospheric pressure and temperature values at varying altitude, and

setting the blowout threshold evaporative efficiency at a fixed 75%, one obtains the

theoretical evaporation-controlled limit curve as is shown in Figure 2.6. Note that the

Chapter 2

19

value of the proportionality constant was arbitrarily chosen for illustration purpose so

that the limit contour passed through Mach 0.87 at an altitude of 10 km.

The reaction-controlled combustion efficiency was based on Equation 2.18 and the

assumption that for an overall global reaction scheme the fuel and oxygen

concentrations could be regarded as referring to the initial concentrations 𝑥𝐹 = 𝜙

𝜙+85.7

and 𝑥𝑂 = 18

𝜙+85.7. With 𝜙 < 1 both 𝑥𝑂 and the denominator of 𝑥𝐹 are nominally

constant and can be absorbed into the overall proportionality. Equation 2.18 therefore

simplifies to 2.30 and employing exponents of 𝑖 and 𝑗 corresponding to the values

recommended by Lefebvre [24], the efficiency expression can be further reduced to

2.31. This was employed to produce the theoretical reaction-controlled extinction limit

contained in Figure 2.6. Again, the overall proportionally constant was arbitrarily chosen

for illustrative purpose only.

𝜂𝜃𝜙�̇�𝐴 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌𝑖+𝑗 𝜙𝑖 [2.30]

𝜂𝜃 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌1.75 �̇�𝐴⁄ [2.31]

The broken curves parallel to the evaporation and reaction-controlled curves are

illustrative of the effect that variations in the proportionally constants have, as would be

expected for different fuels with different evaporation and reaction behaviour.

A number of conclusions were drawn from the theoretical treatment and resultant

Figure 2.6. The mixing-controlled limit is independent of the air flow and air density (to a

first order of magnitude analysis), which discounts it from playing a role in the

discrepancy that Rolls-Royce detected at high altitude and high subsonic Mach numbers.

This is further reinforced by the fact that fuel properties do not directly influence the

parameters contained in the equation governing mixing behaviour (2.11).

The negative gradient of both the evaporation and reaction-controlled limits suggest

that either of these aspects could provide a candidate explanation for the fuel-related

variance that was detected by Rolls-Royce. The fact that fuel properties could

conceivably cause a significant distinction in the manifestation of either of these limiting

efficiency boundaries does not contribute to determining the more probable candidate.

The theoretical treatment is supported by the Honeywell Aerospace test results which

detected very little distinction between the different test fuels. These tests were


20

conducted at 25000 ft. (7.62 km) where the flight envelope is not expected to be

governed by fuel properties.

Figure 2.6: Theoretical thrust limit, constant evaporation- and reaction-controlled combustion efficiency traces in the altitude-Mach number domain

Chapter 4 revisits the efficiency limit treatments in terms of fluid-mechanic and

chemical timescales to examine their relevance under the different combustor

conditions and with different fuels.

2.3 Hypotheses and project plan

The preceding preliminary theoretical treatment identified the root cause of the altitude

blowout differences by the Rolls-Royce report as being determined by either the

evaporation or the reaction-controlled combustion efficiency but not by mixing

efficiency. The following three hypotheses were therefore formulated to define the

scope of the thesis:

Hypothesis 1: A range of aviation jet fuels could be blended according to a design-of-

experiment strategy to exhibit independent variations in properties relating to

evaporation and reaction behaviour whilst still meeting the legislated physical fuel

specifications.

Chapter 2

21

Hypothesis 2: The range of variations contemplated in Hypothesis 1 could have a

detectable and significant influence on the simulated high altitude extinction behaviour

in a representative aviation gas turbine combustor. (Note that this is not a legislated

criterion)

Hypothesis 3: The relevant properties described in Hypothesis 1 can be scientifically

quantified in an appropriate test apparatus and can be correlated to the high-altitude

extinction behaviour described in Hypothesis 2.

The thesis was structured to address these hypotheses primarily through theoretical

analyses of targeted experimental programmes. A schematic summary of the thesis

structure is presented graphically in Figure 2.7.

Figure 2.7: Schematic representation of thesis structure with relevant section numbers

provided in brackets.


22

The evaporation and chemical reaction properties of the test fuels were assessed

against theoretical models prior to being incorporated in the evaluation of the

experimental results. Theoretical models were presented of both evaporation and

reaction limited combustion. The theoretical models, fuel properties and results from

model-combustor experiments were combined to interrogate evaporation and reaction-

controlled extinction behaviour and address hypotheses 1 and 3. By combining these

results with a theoretical analysis of higher altitude and Mach number operation,

Hypothesis 3 was addressed and an explanation was provided for the results of the FSJF

certification tests.

In the following sections a selection of relevant literature is reviewed prior to the

theoretical assessment being expanded in support of the experimental design and

evaluation of test results.

23

3 Literature Review

The combustion performance of practical aircraft combustors is known to be primarily

influenced more by physical processes of heat transfer, mass transfer, thermodynamics,

gas dynamics and fluid dynamics than by chemical processes. While chemical reaction

rates are acknowledged as being of great importance they are largely disregarded in

high temperature flames due to being considered to be relatively rapid and not rate

controlling. Emphasis is placed on large scale mixing and the interdiffusion of fuel and

air as the rate controlling steps. Chemical processes are, however, acknowledged to be

of particular importance for the formation of pollutant emissions and for the lean light-

off and lean blowout (or weak extinction) limits at high flight altitudes [4, 5].

The difficulty in investigating fuel property effects on gas turbine combustion is evident

throughout literature. The interrelated nature of various fuel properties precludes the

classical experimental research approach of quantifying the effects of varying a single

independent parameter while maintaining all others relatively constant. The fact that

different hardware designs have been shown to respond differently to fuel property

variations is a further confounding factor. One notable example can be found in the

response of liner temperature of combustors with either lean or rich primary

combustion zones to carbon/hydrogen (C/H) ratio variations. Heat transfer to the liner

wall is primarily due to radiation in the case of rich primary zone combustion and is thus

governed by flame emissivity, which in turn is influenced by the C/H ratio. In a lean

primary combustion zone heat transfer to the liner wall occurs primarily due to forced

convection which is relatively insensitive to C/H ratio changes with the gas temperature

being dominant [34].

In many instances engine-hardware-oriented studies that focused on investigating the

influences of gas turbine fuel properties on combustion analysed the net effect of fuel

variations on a specific combustor parameter. In some instances attempts were made to

isolate and categorise fuel property effects.

Venkataramani isolated fuel volatility effects from atomisation effects during an

investigation of fuel property influences on altitude relight performance [35]. The study

focused only on separating the influences of physical fuel properties, while chemical

property differences were not investigated. Four test fuels were employed (JP-4, Jet A,

Jet A/2040 solvent blend, and Diesel 2) covering a wide range of volatilities. In order to

Literature Review

24

eliminate fuel specific effects on atomisation, injection equipment was used that

maintained equivalent SMD values (50 ± 10 μm) for all test fuels at equal design flow

rates. The results revealed, inter alia, that fuel volatility assumed a secondary role in

initial (first cup) light-off. Decreased volatility resulted in slightly poorer blowout

performance but full-propagation and first cup blowout were found to be independent

of fuel volatility. In spite of the previously mentioned acknowledgement that chemical

reaction time scales can be significant under threshold conditions such as encountered

during altitude relight, the potential influence of the fuels exhibiting different chemical

reaction rates was not taken into account. Airblast atomisers were shown to exhibit

poorer ignition performance and a stronger volatility dependence than pressure

atomisers.

In another investigation aimed at anticipating the combustion performance effects of

future fuel formulations, Lefebvre analysed a large body of data from studies conducted

by the USAF, Army, Navy and NASA [34]. The matrix of thirteen test fuels included JP-4,

JP-8, five blends each of JP-4 and JP-8, as well as a No. 2 diesel which provided three

levels of hydrogen content: 12, 13 and 14 percent by mass. The results from the studies

were analysed to determine how combustion efficiency, lean blowout limits, ignition

performance, liner wall temperature, emissions and pattern factor were influenced by

the differences in fuel formulation.

Hydrogen content and/or aromatic content exhibited a significant influence on flame

radiance, liner wall temperature and smoke emissions. Physical fuel properties that

influenced atomisation quality and evaporation rates were shown to affect lean

blowout, ignition, combustion efficiency and CO emissions while liner wall temperature,

NOx and smoke emissions were not influenced by fuel physical properties. Combustion

efficiency, lean blowout, ignition and CO and NOx emissions exhibited a small

dependence on fuel chemistry differences. This was attributed to the influence of slight

variations in calorific value on combustion temperature. The physical fuel properties had

an appreciable effect on pattern factor at low power conditions but this effect

decreased to be very small at high power settings where pattern factor effects on vane

life are typically most significant. The combustor pattern factor was not influenced by

fuel chemistry. It should be noted that in this study the carbon to hydrogen (C/H) ratio

was taken as representative of the fuel chemistry differences and no metric

Chapter 3

25

representative of chemical reactivity such as ignition delay or laminar flame speed (LFS),

was measured or reported.

Rao and Lefebvre [36] concluded that the presence of sufficient fuel vapour in the

ignition zone is the sole determining criterion for ignition of heterogeneous mixtures of

fuel droplets in air. Ignition will ensue automatically if sufficient thermal energy is

created by the passage of the spark to produce the required amount of fuel vapour. The

basis of this argument is the understanding that, over a wide range of operating

conditions, the chemical reaction time is very short relative to the time required for

producing an adequate amount of fuel vapour in the ignition zone. This is, however, not

necessarily the case under threshold conditions where the reaction timescales can be

impaired.

Ballal and Lefebvre [37] proposed theoretical models for determining the minimum

ignition energy required and the quench distance in liquid fuel sprays. Quench distance

is defined as the critical spark-kernel size at which the rate of heat loss at the kernel

surface is equal to the rate of heat release throughout the kernel volume, due to the

instantaneous combustion of fuel vapour. The kernel must attain this size in order for

combustion to propagate unaided and the minimum ignition energy is defined as the

amount of energy required from an external source to attain the quenching distance [4].

In a subsequent study the model was extended to include the presence of fuel vapour in

the mixture entering the ignition zone and the influence of finite chemical reaction

rates, which are known to be significant for well atomised fuels at low pressures and low

equivalence ratios [38]. This yielded Equation 3.1 for calculating quenching distance (dq)

for all conditions likely to be encountered in practical combustion systems. A good fit

was obtained with experimental data over SMD values ranging from 40 to 150 μm.

𝑑𝑞 = [𝜌𝐹𝐷32

2

𝜌𝐴𝜙𝑙𝑛(1+𝐵𝑠𝑡)+ (

10𝛼

𝑆𝐿)2

]0.5

[3.1]

The first term of the root sum square equation deals with fuel evaporation. Fuel and air

densities are represented by ρF and ρA respectively, Sauter mean diameter by D32,

equivalence ratio by ɸ, and the stoichiometric mass transfer number by Bst. The second

term introduces a measure of a fuel chemistry influence with α representing the

thermal diffusivity and SL the laminar flame speed of the fuel. By dividing the thermal

diffusivity by the laminar flame speed, the second term effectively reduces to the

Literature Review

26

product of the specific heat at constant pressure (Cp), density (ρ) and laminar flame

thickness ( L ). The influence of the chemical reaction rate is thus reflected in terms of

the laminar flame thickness. The importance of chemical reaction timescales in the

combustion of well atomised fuels at low pressures and lean equivalence ratios as well

as under threshold combustion conditions such as encountered during altitude relight

and blowout is therefore acknowledged to some degree.

Increased interest in alternative aviation fuel formulations, and the related opportunity

to influence chemical reaction rates to a greater degree than traditionally possible in the

case of petroleum-derived jet fuel, has prompted investigation into the influence of

physical and chemical fuel properties on gas turbine combustion behaviour. A number

of recent studies have reported ignition and extinction behaviour being influenced by

fuel formulation to varying degrees. These studies employed experimental setups

ranging from laboratory scale combustors to full scale gas turbines and test fuels

ranging from single component fuels to full boiling range jet fuel formulations. Findings

ranged from reports of appreciable differences to relative insensitivity to fuel

formulation [6 - 13].

Fyffe et al. conducted an altitude ignition and extinction evaluation of five synthetic

paraffinic kerosene (SPK) fuel formulations at the Rolls-Royce plc. sub-atmospheric

altitude ignition facility in Derby, UK [8]. A multi-sector representation of an advanced

gas turbine combustor and fuel injector was used to evaluate threshold combustion at

two test conditions that were considered to be representative of approximate altitude

conditions of 25000 to 30000 ft. (7.62 to 9.14 km). The five test fuels were all GTL-

derived and were compared against a representative crude-derived Jet A-1. The results

were interpreted against three compositional variables that were chosen to define the

fuel matrix: carbon number distribution (narrow vs. wide), iso/normal paraffin ratio and

total paraffinic content. Unfortunately other fuel properties were not reported which

hinders the subsequent critical assessment of the results. Some differences were

reported in ignition behaviour but extinction differences were considered to be within

the experimental scatter. The study associated low iso/normal paraffin ratios with

improved ignition behaviour. This trend is known very well in diesel and gasoline engine

fields where the extremities of the respective cetane and octane rating scales are each

defined by a normal and iso-paraffinic reference fuel. The results were not subjected to

a theoretical assessment and fundamental reasons for the observed behaviour were not

Chapter 3

27

offered but it did support the possibility for fuel formulation differences impacting

threshold combustion.

Burger et al. presented results from a lean blowout (LBO) study of 16 different fuel

formulations in an in-house developed combustor that was based on the primary zone

of an Alison T63. A simplex pressure atomiser nozzle was used in the evaluation. The

authors reported blowout limits correlating with volatility and significantly reported a

(weak) negative correlation between derived cetane number (DCN) and LBO [7].

A follow-on paper evaluated eight test fuels from the same matrix in a representative

aero-engine combustor sector equipped with an airblast atomiser [9]. Significant

variability in the results prevented more definitive conclusions being drawn from this

study. It is notable that the shape of the extinction limit curves suggests that the tests

were conducted in a regime that was primarily influenced by evaporation efficiency. It is

significant that under this regime of operation any potential correlation between DCN

and LBO was masked. This topic is revisited in more detail under Section 4.3.

Rock et al. reported on an experimental study of lean blowout in a swirl-stabilised

combustor equipped with both pressure and airblast atomisers [10]. The study’s key

motivation was similar to this thesis, to investigate the relative roles of physical

properties and chemical kinetic properties on lean blowout behaviour. The test fuel

matrix encompassed eight fuels including three crude-derived jet fuels (Jet-A, JP-5, and

JP-8) and five formulations that spanned a range of physical and chemical properties.

The pressure atomiser blowout tests revealed strong correlations with fuel physical

properties, particularly boiling point temperature. Fuels that were less easily atomised

and vaporised were harder to blowout. This is in accordance with the concept that

delaying atomisation and or vaporisation delays the level of premixing that drives very

lean global fuel-air ratios. It creates localised regions of higher fuel-air ratios and

elevated flame temperatures which supports combustion. The authors proposed that

this behaviour was dependent on the pre-heat temperature being above the fuel flash

point and that the reverse behaviour is exhibited when the pre-heat temperature is

lower than the flash point which would be in agreement with the work by

Burger et al. [7]. It also agrees with the statement by Lefebvre that pressure atomisers

typically produce lower fuel/air mixing prior to combustion and have superior blowout

performance [39]. With the exception of one fuel, the pressure atomiser results also

reported good correlations between blowout and C/H ratio, iso-paraffinic content and

Literature Review

28

fuel smoke point. The authors ascribed this to a strong co-correlation with the more

significant physical parameter T90.

In contrast to the pressure atomiser behaviour, Rock et al. reported no strong

correlation between the blowout behaviour recorded with the airblast atomiser and fuel

physical properties. The strongest correlation that they recorded was with the iso-

paraffinic content of the test fuels which the authors ascribed to being driven by the

chemical kinetic properties of the fuel. The authors indicated that no clear correlation

was found between the airblast blowout results and cetane number which was

supported by substantial scatter in the figures presented. However, closer inspection of

the correlation matrix revealed significant negative correlations with LBO for both

atomiser options. This was further supported by a strong negative correlation between

iso-paraffinic content and cetane numbers. This lacuna within the paper is significant in

view of the contradictory observations highlighted in the earlier works. It is especially

relevant to the present research initiative and provides a motivation for a careful study

of the issue of a possible correlation between the cetane rating of the fuel and LBO.

Grohmann et al. investigated lean blowout differences using the same gas turbine

model-combustor that was employed in this thesis [11]. In an attempt to minimise the

complexity inherent in multi-component jet fuel, they selected six single-component

model fuels to represent normal, branched, cyclic and aromatic hydrocarbons. The

paper reported atomisation, vaporisation, chemical kinetic properties and lean blowout

results for three of these fuels: n-hexane, n-dodecane and iso-octane. The authors

observed that both physical and chemical fuel properties influenced the measured LBO

behaviour. Differences in the LBO behaviour of n-hexane and iso-octane were ascribed

to chemical kinetic nature of the fuels and in agreement with normal alkanes exhibiting

shorter ignition delays than branched alkanes (of the same carbon number). The two

fuels exhibited very similar atomisation and vaporisation behaviour.

By comparison n-hexane and n-dodecane revealed comparable chemical kinetic

characteristics, but very different atomisation and vaporisation behaviour. At high air

pre-heat temperatures the two fuels reported similar LBO results, but at lower pre-heat

the larger droplets and lower evaporation rate of n-dodecane was credited with

extending LBO limits. The authors proposed a similar explanation to that offered by

Rock et al. [10] but interestingly in this instance the atomiser was a prefilming airblast

unit.

Chapter 3

29

Bluff-body flames have been studied extensively, but a lack of consensus about the

physical and chemical processes that lead to blowout motivated a study by Heulskamp

et al. which utilised a large data set of experimental and literature LBO results for bluff-

body stabilised flames to develop a correlation for predicting lean blowout [40]. They

confirmed a dependence on the Damköhler number with fuel variation playing a

significant role. Damköhler number is commonly used in chemical engineering as an

expression of the ratio of chemical and fluid-mechanic timescales. Most researchers

agree that LBO is governed by it, but there have been different proposals about the

definition of the critical timescales. Research such as reported by Shanbhogue et al.

advocate the impact of ignition delay while Radhakrishnan et al. and Kariuki et al.

presented successful correlations with flame speed which argued that blowout is not

governed by ignition timescales [41, 42, 43].

Heulskamp et al. thus endeavoured to provide clarity about critical parameters and their

significance to physical and chemical processes. They concluded that pressure,

temperature and C/H ratio of the fuel impacted the reactivity of the mixture and

contributed to the chemical timescale within the Damköhler number. It was confirmed

that blowout was not solely dependent on the fluid mechanic timescales and LFS was

discredited as a blowout predictor. Ignition delay was found to be an excellent

representation of the chemical timescale in the experimental data set as evidenced by

Figure 3.1 where D/U refers to the fluid-mechanic timescale (U is the lip velocity and D is

the characteristic diameter of the bluff body).

Figure 3.1: Ignition delay dataset correlation results with D/U and IDT as factors

(R2 = 0.918) (from Heulskamp et al. [40])

Literature Review

30

The authors also advocated the influence of cetane number and C/H ratio on blowout

behaviour due to the covariance of these properties with ignition delay. This is in

agreement with Colket et al. who reported fuels with higher cetane numbers and higher

C/H ratios blowing out at lower equivalence ratios as shown in Figure 3.2 [44].

Figure 3.2: Derived cetane number against equivalence ratio at LBO

(from Colket et al. [44])

The work by Won et al. [45] also supported the relevance of cetane number to gas

turbine combustion and extinction. Their study of the pre-vaporised global combustion

behaviour of a selection of petroleum-derived and alternative jet fuels yielded a

‘‘combustion property target (CPT) index” that was developed through regression

analysis of results from three experiments. The CPT index was proposed as a fuel

screening tool for assessing fully pre-vaporized, kinetically coupled behaviours of

potential alternative jet fuel formulations. Derived cetane number (DCN) was included

in the analysis as a measure of autoignition propensity. DCN is calculated directly from

ignition delay measurements in an ignition quality tester (IQT™) which will be discussed

in greater detail under the experimental design in Chapter 5. The resultant CPT index for

extinction employed DCN, heat of combustion, and molecular weight and the CPT index

for low temperature heat release (LTHR) rate was purely based on DCN. Figure 3.3

reproduces a figure reporting the results of the linear regression of the diffusion flame

Chapter 3

31

extinction at a fixed mass fuel fraction (Yf = 0.4). The typical range of JP-8 fuels is

demarcated by the dashed rectangle.

Figure 3.3: CPT index correlation for diffusion flame extinction at Yf = 0.4

(from Won et al. [45])

Figure 3.4 provides a comparison of DCN and the measured LTHR rate as represented by

the maximum H2O mole fraction in the temperature range between 550 and 700 K.

The authors made the pertinent comment about DCN in the context of LTHR that most

of the alternative jet fuels and their blends with petroleum-derived jet fuels were

outside of the variability of JP-8. In light of the approval of 50:50 blends of SPK and

crude-derived fuel they made the point that low temperature reactivity might not be a

critical combustion property for existing gas turbines. However, they highlighted that

advanced engine designs are focused on increasing pressure ratios and combustor

pressures significantly in order to achieve greater efficiency. Under these conditions low

temperature reactivity is expected to play a more significant role under threshold

combustion conditions such as LBO and altitude relight.

Literature Review

32

Figure 3.4: Comparison of DCN to measured LTHR (from Won et al. [45])

Evidence from literature clearly supports the importance of considering both physical

and chemical fuel properties in the context of threshold combustion and high altitude

blowout extinction. There are, however, inconsistencies and omissions in said literature

which necessitate further exploration.

One example is the contradicting conclusions regarding the impact of cetane and LFS

which require clarification. Associated with this; it has been demonstrated that the

comprehensive characterisation of test fuels making use of appropriate metrics is vital.

Other areas that have exposed difficulties included the use of test fuel matrices that are

both representative of real fuels and also sufficiently differentiated to enable

meaningful correlation analyses. The use of technically relevant combustor architecture

and injection equipment is another area that has been shown to impact the conclusions

drawn by different studies.

A very comprehensive, multi-disciplinary, three-year research initiative known as the

National Jet Fuels Combustion Program (NJFCP) is currently in progress in the United

States of America. Its aim is to develop an experimental and analytical capability to

facilitate OEM’s evaluation of fuel physical and chemical properties on engine

operability and to streamline the ASTM fuels approval process. Ignition and lean blow-

Chapter 3

33

out and the influence of flight altitude are some of the principal areas of focus for the

program. Aspects of the literature cited in this chapter are associated with the first fruits

of this program and it is expected that a substantial volume of relevant literature will

continue to emerge from this program during 2017 and beyond.

34

4 Theoretical Assessment

In the following chapter the preliminary theoretical assessment captured in Chapter 2 is

expanded. The handling of combustion efficiency in a chemical kinetics driven system is

extended. The influence of spray and evaporation on extinction limits is discussed and

the significance of the different combustion efficiency definitions is investigated.

4.1 Burning velocity model

It was mentioned that there is more than one approach to defining combustion

efficiency in a system where chemical kinetics are considered to be rate-controlling and

evaporation and mixing rates are assumed to be infinitely fast [46, 47, 48]. The

preliminary theoretical approach in Chapter 2 was based on the well-stirred reactor

model. In the following section the alternative burning velocity model is discussed

followed by an expansion of the stirred reactor model.

The burning velocity model proposes a combustion zone that is analogous to the flame

brush on a Bunsen burner. All the fuel that burns is assumed to burn completely and

combustion inefficiency is attributed to some of the mixture passing through the

combustion zone without being entrained by the turbulent flame front. The combustion

efficiency is then defined as:

𝜂𝜃 = (ℎ𝑒𝑎𝑡 𝑟𝑒𝑙𝑒𝑎𝑠𝑒𝑑 𝑖𝑛 𝑐𝑜𝑚𝑏𝑢𝑠𝑡𝑖𝑜𝑛)

(ℎ𝑒𝑎𝑡 𝑎𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 𝑖𝑛 𝑓𝑢𝑒𝑙)⁄

= (𝜌𝑔𝐴𝑓𝑆𝑇𝑐𝑝∆𝑇

(𝑞�̇�𝐴𝐻)⁄ [4.1]

where ST is the turbulent burning velocity. By definition 𝑐𝑝∆𝑇 = 𝑞𝐻 and with the flame

area (Af) assumed to be proportional to the combustor reference area (Aref) 4.1

simplifies to

𝜂𝜃 = 𝜌𝐴𝑓𝑆𝑇

�̇�𝐴⁄ ∝

𝑆𝑇𝑈𝑟𝑒𝑓⁄ [4.2]

In his studies of Bunsen burner flames Damköhler [49] concluded that the effect of

turbulence on burning velocity depends on the scale of the turbulence relative to the

flame front thickness. He derived the following expression for turbulent burning velocity

Chapter 4

35

where St and SL are the turbulent (small scale) and laminar flame speed respectively, 𝜖 is

the turbulent exchange coefficient, 𝜈 the kinematic viscosity, Re the Reynolds number

and k a constant:

𝑆𝑡 = 𝑆𝐿√𝜖

𝜈 = 𝑆𝐿𝑘√𝑅𝑒 [4.3]

Greenhough and Lefebvre provided a relation between large scale and small scale

turbulence by amending Karlovitz’s expression for large scale turbulence [23]. The

combustion zone is visualised as a region in which small volumina of gas are produced

by strong turbulence. These volumina of gas propagate at a burning velocity that is

governed by the small scale turbulence and the root mean square value of the

fluctuating velocity, �̅�.

𝑆𝑇 = √(2𝑆𝑡�̅�) [4.4]

Combining 4.3 and 4.4 and substituting into 4.2 yields:

𝜂𝜃2 =

2𝑘𝜌2𝐴𝑓2𝑆𝐿�̅� 𝑅𝑒

0.5

�̇�𝐴2⁄ [4.5]

with Reynolds number defined as:

𝑅𝑒 = v𝐷𝜌

𝜇=

�̇�𝐴𝐷

𝐴𝜇 [4.6]

They assumed viscosity to vary little with pressure and thus expressed it as a constant

multiplied by a function of temperature

𝜇 = 𝑘12 𝑇0.75 [4.7]

The fluctuating velocity was expressed in terms of liner pressure loss, ∆𝑃𝐿, where the

fraction of the pressure loss that promotes useful turbulence is denoted by k2.

�̅� = (2𝑔𝑘2 ∆𝑃𝐿

𝜌)0.5

[4.8]

Further by defining the dynamic head as

𝑑 = 𝜌v2

2𝑔=

�̇�𝐴

2𝑔𝜌𝐴2 [4.9]

Equation 4.5 becomes

𝜂𝜃2 = (

2𝑘

𝑘1) (

𝐴𝑓2𝐷0.5

𝐴1.5) (

𝜌2𝑆𝐿2

�̇�𝐴𝑇0.75 ×

𝑘2∆𝑃𝐿

𝑑)0.5

[4.10]

Theoretical Assessment

36

They related the laminar flame speed to inlet pressure and temperature in an

expression of the form

𝑆𝐿 ∝ 𝑃(𝑛−2)/2𝑓(𝑇) [4.11]

where 𝑛 is the reaction order and 𝑓(𝑇) varies with fuel-air ratio and fuel type. This

resulted in Lefebvre arriving at Equation 4.12 containing an exponential temperature

relationship of empirical origin [24, 25]. The combustor reference area, 𝐴𝑟𝑒𝑓, was

assumed to be proportional to the flame area. 𝐷𝑟𝑒𝑓 is the width of the combustor casing

and 𝑞𝑟𝑒𝑓 is the reference dynamic pressure.

𝜂𝜃 = [𝑃3𝐴𝑟𝑒𝑓(𝑃3𝐷𝑟𝑒𝑓)𝑚𝑒(𝑇3 𝑏⁄ ) �̇�𝐴⁄ ][∆𝑃𝐿 𝑞𝑟𝑒𝑓⁄ ]

0.5𝑚 [4.12]

There is clearly a disconnect between the left hand side being dimensionless and the

right hand side having units of [m/s]. Lefebvre and Halls addressed similar concerns

about dimensional inconsistencies by pointing to the introduction of, for example,

viscosity expressed as a constant multiplied by temperature. The subsequent omission

of the constant introduced dimensions in spite of the formulae essentially being

dimensionless. Assigning values of m = 0.75 and b = 300 resulted in Lefebvre’s simplified

Equation 4.13.

𝜂𝜃 = 𝑓[𝑃31.75𝐷𝑟𝑒𝑓

0.75𝑒(𝑇3 300⁄ ) �̇�𝐴⁄ ] [4.13]

4.2 Stirred Reactor model

The well-stirred reactor model was introduced in Chapter 2 based on a derivation by

Strehlow [28] and appeared to be technically more defensible. The derivation yielded

Equation 2.18 (repeated here for ease of reference).

𝜂𝜃𝜙�̇�𝐴 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌𝑖+𝑗𝑥𝐹

𝑖 𝑥𝑂𝑗

[2.18]

The stoichiometric fuel mole fraction, 𝑥𝑠, was “absorbed” into the proportionality

relation. Note that, by assuming the reactor contents to be perfectly homogenous it is

implied that the chemical reaction between the incoming reactants take place at the

temperature and molecular concentrations of the products. Strehlow omitted a pre-

exponential temperature dependence from the Arrhenius expression (as explained by

Laidler [50]) which Lefebvre included (as T0.5) in his equivalent expression:

𝜂𝜃𝜙�̇�𝐴 ∝ 𝑉𝑇0.5𝑒−𝐸 𝑅𝑇⁄ 𝜌𝑖+𝑗𝑥𝐹

𝑖 𝑥𝑂𝑗

[4.14]

Chapter 4

37

The balanced equation for the lean combustion (as would be applicable to LBO) of a

kerosene fuel with a representative chemical formula C12H24 is as follows:

𝜙𝐶12𝐻24 + 18𝑂2 + 67.7𝑁2 = (1 − 𝜂𝜃)𝜙𝐶12𝐻24 + 12𝜂𝜃𝜙(𝐶𝑂2 +𝐻2𝑂) + 18(1 − 𝜂𝜃𝜙)𝑂2

+67.7𝑁2 [4.15]

The fuel and oxygen concentrations are therefore given by

𝑥𝐹 = (1−𝜂𝜃)𝜙

5𝜂𝜃𝜙+𝜙+85.7 [4.16]

and

𝑥𝑂 = 18(1−𝜂𝜃𝜙)

5𝜂𝜃𝜙+𝜙+85.7 [4.17]

Since 𝜙 < 1, the denominators in 4.16 and 4.17 can be considered to be nominally

constant and absorbed into the proportionality relationship. Substitution into 2.18

yields

𝜂𝜃�̇�𝐴 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌𝑖+𝑗(1 − 𝜂𝜃)

𝑖𝜙𝑖−1(1 − 𝜂𝜃𝜙)𝑗

∴ �̇�𝐴

𝑉𝑃𝑖+𝑗= 𝐾

(1−𝜂𝜃)𝑖𝜙𝑖−1(1−𝜂𝜃𝜙)

𝑗

𝑇𝑖+𝑗𝜂𝜃𝑒−𝐸 𝑅𝑇⁄ [4.18]

Note: This expression agrees with Lefebvre’s initial reaction-rate theory treatment [51],

but when it is recalled three chapters later in the stirred reactor development it is

presented as [24]:

𝜂𝜃 = 𝑓[𝑃32𝑉𝑐𝑒

𝑇3 300⁄ �̇�𝐴⁄ ] [4.19]

No justification is provided for the omitted terms and the change in exponential

expression appears fundamentally flawed. It can, however, be shown that over the

typical temperature range and with an appropriate adjustment of the proportionality

constant 𝑒𝑇 300⁄ ≅ 𝑒−400 𝑇⁄ . The expression 𝑒𝑇 𝑏⁄ was proposed in chemistry literature

during the early 1900s as a chemical reaction rate representation, but it has

subsequently been replaced by the Arrhenius formulation, 𝑒−𝐸 𝑅𝑇⁄ , due to it being

technically more justifiable which makes Lefebvre’s choice to revert to the outmoded

form surprising.

Similarly the expression for rich combustion can be derived from the balanced equation:

𝜙𝐶12𝐻24 + 18𝑂2 + 67.7𝑁2 = (1 − 𝜂𝜃)𝜙𝐶12𝐻24 + 12𝜂𝜃𝜙(𝐶𝑂2 𝑜𝑟 𝐶𝑂) + 12𝜂𝜃𝜙(𝐻2 𝑜𝑟 𝐻2𝑂)

+18(1 − 𝜂𝜃)𝜙𝑂2 + 67.7𝑁2 [4.20]


38

The fuel and oxygen concentrations are therefore given by

𝑥𝐹 = (1−𝜂𝜃)𝜙

5𝜂𝜃𝜙+𝜙+85.7 [4.21]

and

𝑥𝑂 = 18(1−𝜂𝜃)𝜙

5𝜂𝜃𝜙+𝜙+85.7 [4.22]

Which for a second-order reaction then yields:

�̇�𝐴


(1−𝜂𝜃)𝑖+𝑗𝜙𝑖+𝑗−1


Equations 4.18 and 4.23 provide the relationship between the volumetric flow rate

through the combustor and the combustion efficiency (the extent to which combustion

was completed during the available residence time) for lean and rich combustion

respectively. When plotted for a particular stoichiometric ratio and typical values of E/R

and T, the general form (Figure 4.1) compares well with similar figures by Strehlow and

Lefebvre.

Figure 4.1: Volumetric flow rate vs. efficiency in a typical well-stirred reactor

The right-hand segment of the curve with a negative slope represents stable

combustion while the peak represents the maximum flow rate condition beyond which

combustion cannot be sustained. The left-hand segment of the curve is not valid since

Chapter 4

39

combustion will always jump to the higher efficiency right-hand side for volumetric

flows below the peak value. Lefebvre and Strehlow differ in their definition of the exact

blowout point on the diagram. Lefebvre introduces a load line which represents the

heat required to raise the reactants to the reaction temperature. At blowout the limiting

load line is tangential to the flow rate curve with the point of tangency identifying the

associated volumetric flow and combustion efficiency. Strehlow on the other hand

maintains that blowout corresponds with the maximum flow rate value. In either case

blowout always occurs in a combustion efficiency range of 70% to 90% [26, 52].

Some pure hydrocarbons were modelled over a range of equivalence ratios for which

peak values of volumetric flow rates were calculated. The variation in the resultant

blowout curves as presented in Figure 4.2, are purely due to the intrinsic differences in

adiabatic flame temperature (AFT). The adiabatic flame temperature calculations

incorporated CO2 dissociation and water-gas-shift reactions. In addition to the different

flame temperatures, the fuels are also expected to have differences in chemical reaction

rates which would result in different proportionality constants (K) in Equations 4.18 and

4.23. The relative behaviour of the fuels in Figure 4.2 is therefore merely illustrative.

Figure 4.2: Illustrative blowout curves

The relative behaviour of synthetic jet fuel and crude-derived Jet A-1 can be

interrogated in a similar fashion. Bester employed surrogate blends of toluene and n-


40

heptane to calculate AFTs for Jet A-1 and synthetic paraffinic kerosene [53]. The blends

were based on matching carbon to hydrogen ratios. Only the initial reactant and final

products were evaluated, without considering the intermediate reactions that occur in

an actual combustion process. Toluene and n-heptane were used since reliable chemical

data for internal energy, enthalpy, and entropy was available (Chemkin). A similar

approach was followed here to calculate the AFTs of representative Jet A-1 and fully

synthetic jet fuel (FSJF). The carbon to hydrogen ratios were based on two-dimensional

gas chromatographic analyses of representative samples of each fuel. Jet A-1 with a

hydrogen mass content of 13.9% was simulated by a blend of 24.86% (vol) toluene and

75.14% (vol) n-heptane. A blend of 18.25% (vol) toluene and 81.75% (vol) n-heptane

was used to simulate the FSJF with a hydrogen content of 14.48% (mass). Employing

these blends, the resultant adiabatic flame temperatures at constant pressure showed a

difference between the crude-derived Jet A-1 and FSJF of 3 to 7 Kelvin for typical

combustion air-fuel ratios. Although the FSJF would be directionally more sensitive to

blowout due to its marginally lower AFT, the difference is too small to have a detectable

impact on the altitude blowout envelope.

Fuel-specific differences in ignition delay could, however, have a significant impact on

pre-exponential constants (K) which in turn could influence relative blowout behaviour.

There were various views expressed in Chapter 3 regarding the suitability of DCN as an

ignition delay metric for comparing the behaviour of fuels under normal gas turbine

combustion conditions and this will also be explored in greater detail later. At this

preliminary stage DCN was, however, employed as a convenient proxy for

demonstrating the potential influence of different chemical reaction rates. IQT™

measurements could be performed easily for both crude-derived Jet A-1 and FSJF

samples. The ignition delays of 4.5 ms and 6.7 ms that were recorded for Jet A-1 and

FSJF respectively represents a significant 49% increased delay for FSJF. Relative

proportionality constants (K values) were proposed based on these relative ignition

delays. Figure 4.3 illustrates the resultant cumulative effect of AFT and relative

proportionality constant differences for these two fuels. The limiting air mass flow for

FSJF is essentially scaled down by 49% which is possibly a simplistic representation since

the complex combustion reactions were treated as a single reaction step. In practice the

reaction rate is expected to be limited within the chemical reaction pathway by steps

which would not necessarily be second order and both fuels may indeed have similar

Chapter 4

41

intermediate reaction steps. However, the fact remains that the calculations were based

on real DCN ignition delay differences, highlighting an intrinsic difference in their

combustion reaction timescales. It also captures the expected direction of the

difference which agrees with some of the results from the FSJF certification process.

Figure 4.3: Illustrative blowout curves for Jet A-1 and FSJF taking into account AFT

alone and the cumulative effect of AFT and proportionality constant differences.

Well-stirred, homogeneous reactors differ significantly from heterogeneous

combustors, especially in terms of the local stoichiometry and relevant temperature

regimes. The homogeneous character of a stirred reactor, however, makes it particularly

amenable to theoretical analysis. Weiss et al. and Strehlow found that the blowout

behaviour of fuels in a bluff-body stabilised regime correlated well with the lean

blowout behaviour in a spherical stirred reactor. The implication is that the composition

in a recirculation zone behind a bluff-body at its lean-limit holding efficiency is related to

the composition inside a well-stirred reactor operating at its blowout efficiency [54, 55].

4.3 Evaporation-controlled stability limits

The influence of evaporation can be depicted in a similar manner to the reaction-

controlled blowout curves above in order to predict the expected interaction of both


42

evaporation and reaction limits. Recalling Equation 2.6 the evaporation limited

combustion efficiency can be expressed as follows



⁄ [2.6]

The Sauter mean diameter for prefilming airblast atomisers operating with low viscosity

liquids such as kerosene are approximated reliably by Equation 2.28 which simplifies

further to Equation 4.24 for a given burner and fuel.

𝑆𝑀𝐷 ∝ (𝜎


(1 +1

𝐴𝐿𝑅) [2.28]

𝑆𝑀𝐷 ∝ 1

�̇�𝐴(1 +

1

𝐴𝐿𝑅) [4.24]

Substituting 4.24 as the starting diameter into 2.6, ignoring second order effects, and

assuming 𝜌𝐴 and 𝜆𝑒𝑓𝑓 to be constant for a laboratory burner yields Equation 4.25

𝜂𝑒 ∝ �̇�𝐴

(1+ 1

𝐴𝐿𝑅)2 [4.25]

By assuming a constant combustion efficiency limit the volumetric flow rate through the

combustor reduces to a polynomial function of the equivalence ratio (4.26).

�̇�𝐴

𝑉𝑃𝑖+𝑗∝ (1 + 𝜙)2 [4.26]

The effect of both reaction-controlled and evaporation-controlled efficiency limits are

illustrated in Figure 4.4. This figure is of particular interest when considering some of

the comments contained in the Rolls-Royce test report included as an appendix to the

FSJF certification report [13, 14]. The differences in altitude extinction behaviour of the

two fuels were explained with the use of an illustrative schematic that is reproduced in

Figure 4.5. It was noted that the fuels exhibited negligible differences at the lean and

rich extremities of the stability boundaries, but notable differences near the toe of the

curve. In light of the preceding theoretical treatment, the effect of (fuel-related)

differences in evaporation are not expected to be revealed at the toe of the blowout

curve. In fact with increasing volume flow through the combustor, the evaporation

limits are expected to widen and become less of a constraint. The evaporation is only

expected to constrain the toe region where the reaction boundary is relatively impaired.

Evaporation is therefore not expected to cause a difference in the relative shape of the

Chapter 4

43

toe. This would argue against a connection existing between the differences that were

observed in the blowout limits and evaporation-related fuel properties.

Figure 4.4: Illustrative reaction and evaporation limited extinction blowout curves

Figure 4.5: Extinction performance for simulated wind milling operating condition,

schematic representation of the result. (LUFF = Fuelling line) (from Rolls-Royce report [14])


44

4.4 Combustion Efficiencies

Evaporation-, mixing-, and reaction-controlled efficiencies can also be considered in

terms of fluid-mechanic and chemical timescales to gain insight into their relevance

under the different combustor conditions and how they potentially impact combustion

limits. The usefulness of representing combustion efficiency in this way will be seen

later in Chapter 7.

4.4.1 Evaporation-controlled efficiency

Equation 2.6 was used to express the evaporation-controlled combustion efficiency in

terms of the combustor operating conditions (𝜌𝐴 , 𝜆𝑒𝑓𝑓 , �̇�𝐴), physical combustor

dimensions (V), fuel spray characteristics (D0) and fuel properties (D0 , 𝜆𝑒𝑓𝑓).

Rearranging Equation 2.6, the evaporation-controlled combustion efficiency can be

expressed as Equation 4.27.

𝜂𝑒 ∝ (𝜌𝐴𝑉

�̇�𝐴) (

𝜆𝑒𝑓𝑓

𝐷02 ) [4.27]

The first term on the right hand side of the equation is equivalent to the residence time

of the air in the combustion chamber, 𝜏𝑟𝑒𝑠 and, as discussed in Appendix A, the second

term is the inverse of the droplet lifetime or evaporation delay time, 𝜏𝑒. The expression

therefore represents the logical argument that the combustion efficiency amounts to

the ratio of the air residence time in the combustor to the droplet evaporation time.

𝜂𝑒 =𝐴𝑖𝑟 𝑟𝑒𝑠𝑖𝑑𝑒𝑛𝑐𝑒 𝑡𝑖𝑚𝑒

𝐷𝑟𝑜𝑝𝑙𝑒𝑡 𝑒𝑣𝑎𝑝𝑜𝑟𝑎𝑡𝑖𝑜𝑛 𝑡𝑖𝑚𝑒=

𝜏𝑟𝑒𝑠

𝜏𝑒 [4.28]

It is possible to illustrate the blowout limit of a given fuel (and at a chosen combustion

efficiency limit) by expressing both the air residence time and evaporation time in terms

of air mass flow rate. In a laboratory model-combustor operated at constant pressure

and temperature, the fluid-mechanic timescale, or air residence time, becomes a

function of the inverse of the air mass flow rate. Recalling Equation 2.28 for SMDs and

assuming constant hardware dimensions, fuel properties, the relative velocity and

associated effective evaporation constant of finely atomised droplets being relatively

insensitive to the air mass flow rate [20, 56], and with 1 +1

𝐴𝐿𝑅≈ 1 the droplet

evaporation can be approximated as:

𝜏𝑒 = (𝐷02

𝜆𝑒𝑓𝑓) ∝ (

1

𝑚𝐴̇2) [4.29]

Chapter 4

45

Figure 4.6: Illustrative timescales in an evaporation limited system with fixed inlet

pressure and temperature

Figure 4.6 represents an evaporation-controlled system where the limiting combustion

efficiency has been assumed to be at 80%. Combustion can be sustained as long as the

residence time is greater than 80% of the evaporation delay. This would dictate a

minimum air mass flow rate at which blowout occurs and below which combustion is

impossible. At any airflow rates above the blowout point the air residence time would

be sufficient for the evaporation efficiency to be greater than 80% and evaporation

would not play a limiting role. This confirms the conclusion from Section 4.3 and

Figure 4.4 that evaporation limits are expected to constrain combustion at lower air

mass flow rates. Fuel-related differences in properties that effect evaporation would be

relevant under these conditions, but play a much less important role at higher mass flow

rates.

4.4.2 Mixing-controlled efficiency

The combustion efficiency in a mixing-controlled system was expressed by

Equation 2.11.

𝜂𝑚 = 𝑓 ((𝑃3𝐴𝑟𝑒𝑓 �̇�𝐴𝑇30.5⁄ )(∆𝑃𝐿 𝑃3⁄ )0.5) [2.11]


46

∴ 𝜂𝑚 ∝ (𝜌𝐴𝑟𝑒𝑓

�̇�𝐴) (

𝑇3∆𝑃𝐿

𝑃3)0.5

[4.30]

∴ 𝜂𝑚 ∝ (𝜌

�̇�𝐴) (

𝑇3(𝜌𝑣2)

𝑃3)0.5

[4.31]

∴ 𝜂𝑚 ∝ (𝜌𝐴𝑟𝑒𝑓𝑣

�̇�𝐴) [4.32]

Assuming a fixed dimensionless ratio to exist between the area of the liner air holes and

the combustor reference area (Aref), the numerator is also equal to the air mass flow

rate. This indicates that both 𝜏𝑟𝑒𝑠 and 𝜏𝑚𝑖𝑥 are proportional to 1/�̇�𝐴.

Figure 4.7: Illustrative timescales in a mixing limited system with fixed inlet pressure

and temperature

For any given pressure and temperature condition, the efficiency is nominally

independent of air mass flow rate and is essentially dependent only on the ratio of the

area of the liner air holes to the combustor cross-section area.

4.4.3 Reaction-controlled efficiency

Rearranging Equation 2.31 for reaction-controlled combustion efficiency yields

Equation 4.27.

𝜂𝜃 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌1.75 �̇�𝐴⁄ [2.31]

Chapter 4

47

∴ 𝜂𝜃 ∝ (𝜌𝑉

�̇�𝐴) (𝑒−𝐸 𝑅𝑇⁄ 𝑃0.75) [4.27]

The first term on the right hand side of the equation is again equivalent to the residence

time, 𝜏𝑟𝑒𝑠 , while the second term is the inverse of a common expression for the

reaction induction time or so-called ignition delay time, 𝜏𝜃 [57]. The combustion

efficiency is therefore in fact simply being expressed as the air residence time divided by

the combustion induction time which is equivalent to the Damköhler number, Da:

𝜂𝜃 =𝐴𝑖𝑟 𝑟𝑒𝑠𝑖𝑑𝑒𝑛𝑐𝑒 𝑡𝑖𝑚𝑒

𝐶𝑜𝑚𝑏𝑢𝑠𝑡𝑖𝑜𝑛 𝑖𝑛𝑑𝑢𝑐𝑡𝑖𝑜𝑛 𝑡𝑖𝑚𝑒=

𝜏𝑟𝑒𝑠

𝜏𝜃 ≡ 𝐷𝑎 [4.28]

In a laboratory model-combustor operating at constant pressure and temperature the

fluid-mechanic timescale or air residence time becomes a function of the inverse of the

air mass flow rate while the chemical timescale will be constant. Figure 4.8 represents

this diagrammatically for a system where the limiting combustion efficiency has again

been assumed to be at 80%. Combustion can be sustained as long as the residence time

is greater than 80% of the ignition delay which dictates a maximum air mass flow rate at

which blowout occurs and beyond which combustion is impossible.

Figure 4.8: Illustrative timescales in a chemical reaction limited system with fixed inlet

pressure and temperature.


48

The diagrams presented in this section were chosen to illustrate the relative timescale

behaviour in a combustor having fixed inlet pressure and temperature. Obviously one

could consider the relative timescale behaviour under other scenarios such as fixed air

mass flow and varying inlet pressure or temperature. However, it was the principle of

viewing the efficiency as the ratio of two competing timescales that is useful in this

present work.

The subsequent chapter details the experimental approach that was designed to

interrogate the thesis hypotheses by examining how the theoretical evaporation- and

reaction-controlled combustion efficiencies translate to practical combustion behaviour.

49

5 Experimental Design

An experimental programme was designed based on the insights gained from the

theoretical treatment and literature review to provide data to explore how fuel

properties can impact combustion efficiency and extinction under representative gas

turbine combustion conditions. The programme was also intended to provide data for

answering the hypotheses and explaining the LBO differences that were observed

during the fully synthetic jet fuel (FSJF) certification tests. It consisted of the following

main components:

Design of a test fuel matrix

Physical and chemical characterisation the fuel matrix

Evaluation of the test fuel LBO behaviour in a laboratory-scale model-combustor

Evaluation of the test fuel spray behaviour in a laboratory-scale model-combustor

5.1 Test fuel matrix

The theoretical assessment implicitly proposed a research strategy for revealing the root

cause if it were possible to evaluate a number of fuels that exhibit independently varied

properties relating to evaporation and reaction behaviour. It is, however, virtually

impossible to change fuel properties entirely orthogonally which complicates the

selection of appropriate test fuels and the interpretation of test results. The design of

the test fuel matrix was influenced by a combination of the desire to include fuels that

were representative of those employed during the FSJF certification process and results

from previous laboratory-scale extinction and ignition work [7, 58] which included both

full boiling range and single component test fuels. The primary goal was to investigate

the relative influence of fuel properties on evaporation rate and reaction rate

independently in a combustion environment. It was therefore important to differentiate

key properties to as large a degree as possible, while still employing fuels that

resembled current commercial jet fuels. The resultant matrix entailed a selection of

eight fuels that were nominally representative of either crude-derived or viable

alternative jet fuel formulations. It included fuels that complied fully with the Jet A-1

specification such as crude-derived and Fischer–Tropsch (FT) derived synthetic

commercial fuels as well as fuels that were formulated to intentionally breach the

specification on selected properties that were of interest. Table 5.1 reflects the test fuel

Experimental design

50

matrix with a selection of crude-derived Jet A-1, synthetic paraffinic kerosene, linear

paraffinic solvents, aromatic solvents and pure compounds. While the matrix included

some synthetic jet fuel formulations, the primary focus was on fuel properties

regardless of the production pathways employed to produce the fuels.

Table 5.1: Test fuel matrix

CJF Crude-derived Jet A-1

CJF/D Jet A-1 + 50% n-dodecane

SJF1 FSJF (certification)

SJF2 FSJF (commercial)

SJF2+ FSJF (commercial) + 1.5% HCPP

LTSK Experimental GTL kerosene

HTSK Synthetic paraffinic kerosene (SPK)

HN Heavy naphtha refinery stream

CJF was a commercial crude-derived Jet A-1 while CJF/D was a 50:50 blend of CJF and n-

dodecane. SJF1 was formulated to be representative of the FSJF that was used during

the 2006 certification process and the associated test program, while SJF2 was

representative of a FSJF formulation that would comply with the distillation gradient

specification that was introduced for FSJF as part of DEFSTAN 91-91 [59] in 2008. SJF2+

was a blend of 98.5% SJF2 and 1.5% hydrogenated cat-poly petrol (HCPP) which was

added to investigate the influence of flash point. HTSK was a Synthetic paraffinic

kerosene stream that is typically employed in both fully and semi-synthetic jet fuel. HN

was a heavy naphtha refinery stream and LTSK a gas-to-liquids (GTL) kerosene that were

both included to assist in evaluating non-orthogonal changes in the various fuel

properties. LTSK was of further interest due to the fact that it nominally resembled a

fuel that could also be produced via a renewable synthetic jet fuel manufacturing

process. The influence of the LTSK’s largely linear paraffin structure on ignition delay,

cetane, and LBO was also of interest in light of the observations that have been

highlighted by literature as discussed in Chapter 3 [8, 10, 11].

Chapter 5

51

5.2 Physical characterisation of test fuel matrix

The test fuel matrix was characterised employing the standard test methods employed

in the aviation industry as specified by ASTM D1655 [60]. While full-specification

analyses were performed, the physical properties that were of primary interest included

density, viscosity, flash point, and the distillation profiles. Concerns have been raised

within the aviation industry regarding the sensitivity of the D86 method that is

stipulated by ASTM D1655 for determining distillation profiles. Therefore, apart from

the conventional D86 method, a GC-based simulated distillation was also conducted

according to ASTM D2887 [61]. Surface tension measurements were performed at 25°C

and 40°C, and refractive indices were determined to be employed by the Phase Doppler

Anemometry (PDA) measurements during the fuel spray characterisation.

5.3 Chemical characterisation of test fuel matrix

In addition to the normal (physical) jet fuel specification tests a number of

measurements were conducted to characterise the chemical composition and reactivity

of the test fuels. Comprehensive two–dimensional gas chromatography (GCxGC), where

two independent separations are applied to a single sample injection, is considered to

be the most accurate means to obtain detailed compositional information for complex

petrochemical mixtures. This technique offers structured separations, high sensitivity

and high peak capacity. Complex fuels are traditionally analysed in the so-called

“normal” GCxGC mode where a non–polar column is used in the first dimension to

perform a boiling-point based separation. A thermal modulator is used to focus the

effluent from the first column onto a polar column which provides separation in the

second dimension based on polarity. For this study the analyses were performed in the

so-called “inverse mode” where the first dimension employs a polar column and the

second dimension employs a non–polar column. This sequence is considered to be less

orthogonal but has been demonstrated to extend the separation space when used to

characterise FT fuels and the aromatic fractions in petroleum middle–distillates resulting

in improved resolution [62]. Chemical structures were elucidated by time–of–flight mass

spectrometry (TOF–MS), and a flame ionisation detector (FID) was used for

quantification. The two-dimensional gas chromatography results provided information

about the test fuel chemical compositions and were used to determine carbon to

Experimental design

52

hydrogen ratios that were required to calculate the stoichiometric fuel/air ratio (FAR)

for each test fuel.

Tests were also conducted to characterise the chemical reactivity of the fuels. An IQT™

combustion bomb (manufactured by AET, Canada) was employed to determine ignition

delay times of the test fuel matrix. This automated test method is used for the

quantitative determination of the ignition characteristics of specifically middle-distillate

fuels and fuel blending components. The IQT™ utilises a constant-volume combustion

chamber that is pre-heated electrically to 843 K. Test fuel is injected directly into the

chamber filled with synthetic air at 2.1 MPa. An empirical calibration formula, as

specified in ASTM D 6890 [63], is used to convert the measured ignition delay to a

derived cetane number. This is a combustion quality indicator used primarily in diesel

engine applications.

Laminar flame speed (LFS) has been suggested as a combustion indicator that could

potentially be appropriate for evaluating gas-turbine performance [42, 43, 64]. A

pressure-based method making use of a spherical bomb was employed to measure the

LFS and Markstein length of the test fuels. Tests were conducted at ambient initial

pressure and two initial temperatures with six different air-fuel ratios spanning the

range from lean to rich. A minimum of six repeat combustion pressure records of

approximately 90 data points were obtained for each test fuel which yielded a database

of over 6000 individual calculations of LFS from which the relevant parameter

coefficients could be calculated using a regression technique. The experimental

hardware and analysis methodology are described in greater detail by Yates et al. [64].

Peak LFS values were calculated at an equivalence ratio of ɸ = 1.075 and reference

conditions of 298.15 K and 101.325 kPa.

Ignition delays were determined by the Institute of Combustion Technology at the

German Aerospace Centre (DLR) in Stuttgart, Germany. A high pressure shock tube was

employed with an internal diameter of 98.2 mm. The 5.18 m long driver section and

11.12 m long driven section were separated by aluminium diaphragms. A turbo-

molecular pump was used to evacuate the driven section to pressures below 10-6 mbar.

Gas mixtures were prepared manometrically in an evacuated stainless steel storage

cylinder at pressures below 10-6 mbar. Four piezo-electric pressure sensors at 20 cm

intervals were used to calculate the incident shock speed. A one-dimensional shock

model was used to calculate the temperature and pressure behind the reflected shock

Chapter 5

53

wave from the measured incident shock speed and the speed attenuation. The resultant

uncertainty in the calculated reflected shock temperature was less than 10 K. Piezo-

electric pressure sensors (PCB® 113A24 and Kistler® 603B) located 1 cm from the end

flange, were used to observe the pressure profiles associated with ignition.

OH*-emission (at 308 nm) and CH*-emission (at 431 nm) were measured using a

photomultiplier and narrow band filter (full width at half maximum - FWHM of 5 nm).

The ignition delay time values were calculated as the time elapsed between the system

being initiated by the reflected shock wave and the maximum OH*-emission being

recorded. Ignition delay times of up to 6.5 ms can be measured by the experimental

setup depending on the temperature. All tests were performed at initial pressures of

1.6 MPa ± 10% with synthetic air (20%v O2 in N2). Reaction mixtures were diluted 1:2,

i.e. one part undiluted mixture and one part N2.

5.4 Gas turbine model-combustor: LBO evaluation

The LBO behaviour of the eight test fuels were evaluated using a laboratory-scale

model-combustor at the German Aerospace Centre (DLR). The combustor was based on

previous spray burner research [65] and was designed to be representative of the

primary zone encountered in real aero-engines. The scaling of the combustor was a

compromise to accommodate practical laboratory constraints while retaining technical

relevance. The resultant rig was designed for thermal powers of approximately 10 kW at

atmospheric pressure conditions.

The core of the combustor consisted of the burner nozzle shown in Figure 5.1. A

common air plenum supplied air to inner and outer air swirlers that produced two co-

axial, co-rotating swirl flows that were separated by a thin annular ring with a sharp

edge. A pressure-swirl atomiser produced a hollow cone of fuel spray through two swirl

channels that was sprayed onto the inside of the annular ring. A fuel film formed on the

surface and was transported to the lip of the ring where it re-atomised. This so-called

prefilming airblast atomisation concept is widely used in current combustion

systems [66].

Experimental design

54

Figure 5.1: Laboratory-scale model burner nozzle (configuration B) [65]

The combustion chamber consisted of four quartz windows that were held in place by

steel posts. The resultant optical access allowed the application of a host of optical and

laser-based diagnostic techniques. A top plate secured the windows and posts. The

19 mm diameter exit port in the top plate exhausted to atmospheric pressure. The

chamber had a cross section of 85 mm by 85 mm and a length of 169 mm which was

sufficient to fully contain combustion throughout the operating conditions that were

investigated. Specific care was taken to ensure cylindrical symmetry of the spray cone

and resultant combustion zone.

The experimental layout of the model-combustor is shown in Figure 5.2. A mass flow

controller (Bronkhorst EL-FLOWselect F-203AV) was used to regulate supply of dry

compressed air to the air plenum. The air temperature was controlled by an electric pre-

heater based on the temperature inside the air plenum measured using a K-type

thermocouple. A sonic nozzle decoupled the combustor acoustically from the air feed

line while meshes were used to homogenise the air flow into the plenum. The test fuel

was supplied from a piston driven stainless steel cylinder which ensured that the

nitrogen driver gas was not in contact with the test fuel. The test fuel was also degassed

prior to testing by applying a mild vacuum to the supply cylinder. A mass flow controller

(Bronkhorst mini CORI-FLOW M14) regulated the fuel supply to the fuel pressure

atomiser via a water-cooled fuel lance. The fuel temperature control was based on the

temperature measured just upstream of the atomiser exit. The development of the test

apparatus is discussed in greater detail by Grohmann et al. [65].

Chapter 5

55

Two burner configurations (A and B) were tested with different swirl numbers. The swirl

number is defined as the ratio of the angular momentum flux, 𝐺𝜃, divided by the axial

momentum flux, 𝐺𝑦, and a characteristic radial distance, R [67, 68].

𝑆 = 𝐺𝜃

𝐺𝑦𝑅 [5.1]

Subsequent to the completion of the LBO tests with burner configuration A,

stereoscopic particle image velocimetry (PIV) measurements were conducted of the

non-reacting and reacting flows during an unrelated test program at DLR. The PIV results

revealed the absence of a strong central recirculation zone as would be present in

typical real aero-engine combustors. This was ascribed to lower geometric swirl with a

centre swirl number of Sc = 0.6. The burner was modified, and the resultant second

configuration (B) exhibited a well-developed central recirculation zone with a

geometrical swirl number of Sc = 1.17 for the centre flow and Sa = 1.22 for the annular

flow.

Figure 5.2: Laboratory-scale model-combustor layout [65]

The LBO tests of configuration A were conducted at two air pre-heat temperatures,

323 K and 413 K, while the tests of configuration B were only conducted at 323 K due to

the limited volume of test fuel available. The fuel temperature was controlled to 303 K

Experimental design

56

for the low air pre-heat tests and to 323 K for the high air pre-heat tests. For each air

pre-heat condition, a series of tests were conducted over range of air flow rates ranging

from 2.2 g/s to 12.9 g/s. The air mass flow rate was kept constant for each test point

operating condition. Combustion was established at a stoichiometry of approximately

ɸ = 0.7, well within the stable combustion operating envelope, following which the fuel

mass flow rate was reduced at a controlled rate of 0.5 (g/h)/s up to the point of

blowout. This corresponded with a ɸ reduction rate of 0.001 s-1 to 0.0001 s-1 depending

on the air mass flow rate. The stoichiometry and air mass flow rate at the point of

extinction was recorded as the LBO point. A minimum of three repeat tests were

conducted at each operating condition. The model-combustor set-up and a

stoichiometric Jet A-1 reference flame are shown in Figures 5.3 and 5.4 respectively.

Figure 5.3: Model-combustor setup Figure 5.4: Reference Jet A-1 flame (ɸ = 1, Tair = 323 K)

5.5 Gas turbine model-combustor: fuel spray evaluation

A three component Phase Doppler Anemometry (PDA) system by Artium Technologies

Inc. (PDI-300 MD) was used to investigate the fuel spray in the model-combustor. The

droplet diameters as well as radial, axial and circumferential velocity components of the

Chapter 5

57

droplets were measured during combustion close to LBO. Three pairs of laser beams

were generated by three diode pumped solid state lasers contained in the two

transmitting units: 532 nm for axial velocity and diameter measurements, and 491 nm

and 561 nm both for radial and circumferential velocity measurements. Negative

velocity measurements were enabled by frequency shifting one beam of each of the

beam pairs using a Bragg cell. The receiving unit collected the signal in a 45° forward

scatter configuration. The focal length of the optics was 350 mm with the beam waist in

the measurement volume of between 150 and 180 µm depending on the 3 wavelengths.

Droplet measurements of between 0.7um and 100um were attainable with the spatial

aperture of 500um that was applied. The characterisation of smaller droplets was

optimised by adjusting the amplifier voltages of the signal receiving photo multipliers

based on local measurement requirements which minimised saturation of the photo

multipliers while still maximising small droplet signal. A refractometer was used to

determine the refractive indices of the eight test fuels at room temperature.

Measurements were conducted at two stoichiometric air-fuel ratios: one at the LBO

stoichiometry plus 5% and another at a stoichiometry of ɸ = 0.6. This allowed both

investigation of the spray behaviour as close to blowout as practical, as well as a

comparison between all fuels at exactly the same stoichiometric air-fuel ratio. For each

stoichiometric set point, measurements were performed at two air flow rates: 2.2 g/s

and 6.5 g/s which allowed evaluation both below and above the flow rate range where a

distinct air flow-related influence was detected in the LBO behaviour as discussed in the

following chapter.

The combustor was mounted on a three-axis traversing stage that was driven by stepper

motors. This allowed the positioning of the measurement volume within the

combustion zone. The coordinate system is shown in Figure 5.2. Measurements were

performed at three distances from the nozzle exit plane: z = 15 mm, z = 25 mm, and

z= 35 mm. At each z-axis location the combustor was traversed along the y-axis in 2 mm

increments. A minimum of 2500 droplet measurements, that were coincident for all

three channels, were recorded at each measurement point. This resulted in diameter

values being based on between 3000 and 135000 readings per position depending on

the position and operating conditions.

58

6 Results

The results of the various test programmes are presented in the following chapter. The

general fuel characterisation is presented first followed by the results of the LBO and

PDA measurements that were conducted in the model-combustor.

6.1 Physical characterisation results

Comprehensive fuel property analysis results are reported in Appendix B. The most

relevant properties are listed below in Table 6.1, and the distillation profiles are shown

in Figure 6.1.

Table 6.1: Test fuel characterisation: key physical properties from specification analysis

Test Fuel CJF CJF/D SJF1 SJF2 SJF2+ LTSK HTSK HN

Flash point [K] 315.5 320.0 315.5 326.5 311.0 311.0 319.5 320.0

T5 [K] 437.3 454.2 436.4 446.4 442.3 427.8 442.9 436.6

T50 [K] 468.7 478.6 445.7 469.6 469.2 441.8 451.0 438.7

T90 [K] 508.1 494.6 468.7 514.5 515.2 461.0 469.4 445.3

T50-T10 [K] 25.3 18.0 9.6 19.7 25.2 12.7 6.8 2.0

T90-T10 [K] 64.7 34.0 32.6 64.6 71.2 31.9 25.2 8.6

Density @ 20°C [kg/l] 0.793 0.773 0.795 0.812 0.810 0.730 0.757 0.756

Viscosity @ 40°C [cSt] 1.20 1.29 1.05 1.46 1.42 0.91 1.16 0.91

Surface tension @ 25°C

[mN/m] 25.68 25.61 24.06 25.97 25.94 23.45 23.85 24.25

6.2 Chemical characterisation results

The results of the comprehensive two–dimensional gas chromatography (GCxGC)

characterisation, the laminar flame speed (LFS) measurements, the derived cetane

numbers (DCN) acquired in the IQT™, and shock tube ignition delays corresponding to

three reference temperatures are summarised in Table 6.2.

Chapter 6

59

Figure 6.1: Test fuel volatility: distillation profiles.

Table 6.2: Test fuel characterisation: GCxGC speciation, DCN, LFS and shock tube ignition delays


normal-paraffins [wgt%] 24.0 60.3 8.2 2.5 2.4 67.9 0 38.1

iso-paraffins [wgt%] 28.3 14.9 65.1 43.6 44.5 31.1 92.0 34.7

cyclic paraffins [wgt%] 20.4 10.6 9.1 14.4 14.3 1.0 7.6 17.6

bicyclic paraffins [wgt%] 4.6 2.3 2.8 20.8 20.4 0 0 1.6

tricyclic paraffins [wgt%] 0.3 0 0 7.5 7.4 0 0 0

alkyl benzenes [wgt%] 15.7 8.2 12.8 2.9 2.9 0 0 7.6

cyclic alkyl benzenes [wgt%] 4.4 2.3 1.8 8.0 7.8 0 0 0.4

naphthalenes [wgt%] 2.3 1.2 0 0.3 0.3 0 0 0

C/H ratio 1.93 2.04 2.02 1.95 1.95 2.20 2.16 2.09

LFS [m.s-1] 0.358 0.353 0.345 0.351 0.350 0.338 0.345 0.352

IDIQT [ms] 4.50 3.45 6.68 7.06 7.14 3.24 7.14 3.88

DCN 45.9 58.5 32.4 30.9 30.6 62.0 30.6 52.6

ID1250 K (1000/T = 0.8) [ms] 0.29 0.31 0.34 0.26 0.26 0.22 0.32 0.24

ID1000 K (1000/T = 1.0) [ms] 2.52 2.61 3.05 2.34 2.31 1.94 2.92 2.07

ID833 K (1000/T = 1.2) [ms] 13.25 9.08 22.90 17.07 15.09 10.08 21.11 9.94

Results

60

The IQT™-measured ignition delays (IDIQT) were converted to DCN values using the

empirical calibration described by Equation 6.1 [69].

𝐷𝐶𝑁 = 83.99 (𝐼𝐷𝐼𝑄𝑇 − 1.512)−0.658

+ 3.547 [6.1]

The constant 1.512 ms offset that was applied to all the measured ignition delays

accounted for the evaporation and mixing delays that are intrinsic to the heterogeneous

injection in the IQT™ bomb.

While the physical characterisation, compositional analyses, laminar flame speed, and

IQT™ ignition delay results could be used without significant processing, the data from

the shock tube ignition delay experiments had to be analysed further. The data points

spanned both the high-temperature and low-temperature regimes with a number but

not all of the fuels recording delays into the negative temperature coefficient (NTC)

region. Yates et al [70] proposed a convenient method of describing the ignition delays

in each of the distinct temperature regimes using generic Arrhenius functions having the

form:

𝜏𝑖 = 𝐴𝑖𝑝𝑛𝑖𝑒

𝐵𝑖𝑇 [6.2]

Where 𝜏𝑖 represents the ignition delay, A, n and B, are constants and p and T are the

pressure and temperature respectively. They treated the two sequential stages in the

low-temperature regime (low temperature and NTC) as the arithmetic sum of the

individual delays, 𝜏1 + 𝜏2. The high temperature delay, 𝜏3, represented an alternative

competing pathway and the overall ignition delay for both low and high-temperature

was expressed by Equation 6.3.

𝜏𝑜𝑣𝑒𝑟𝑎𝑙𝑙 = ((𝜏1 + 𝜏2)−1 + (𝜏3)

−1)−1 [6.3]

This approach was applied to the shock tube results to smooth over the experimental

noise and intrinsic test-to-test variation in measurement conditions in order to facilitate

direct comparisons between the various fuels. It also facilitated the inclusion of the

ignition delay results in correlation analysis of the fuel properties and LBO test results.

The values of the coefficients A and B for each of the stages were determined by means

of a regression analysis. The pressure dependence was ignored (𝑛𝑖 = 0) due to all tests

being conducted at a common pressure. The shock tube data set did not contain results

that extended into the linear portion of the low-temperature region for all the test fuels,

which essentially reduced the three-Arrhenius to a two-Arrhenius treatment without

Chapter 6

61

influencing the model fit or accuracy in the NTC or high-temperature regions. An

example of the measured results and modelled ignition delay for one of the fuels (test

fuel CJF) is presented in Figure 6.2. The results, modelled ignition delay curves, and

model coefficients for the full fuel matrix are presented in Appendix C. The modelled

ignition delay results for the full matrix are presented in Figure 6.3 while values

calculated for each fuel at 833 K, 1000 k and 1250 K (corresponding with 1000/T values

of 1.2, 1.0 and 0.8 respectively) are included in Table 6.2.

Figure 6.2: Test fuel CJF: Shock tube measured (markers) and modelled (curve)

ignition delay results based on eq. 6.3.

Results

62

Figure 6.3: Modelled ignition delay results for the full tests fuel matrix

The differences in the absolute values of the ignition delays measured in the

homogeneous shock tube and heterogeneous IQT™ are a manifestation of the inherent

differences in operating conditions. It can, however, be expected that fuels with similar

evaporative behaviour exhibit a correlation between IQT™ and shock tube ignition

delays under similar pressure and temperature conditions by virtue of the fuel-to-fuel

differences in the IQT™ ignition delays being reduced primarily to differences in

chemistry. The shock tube measurements in this study were conducted at 16 bar which

did not allow direct comparison with the delays that were measured at 21 bar in the

IQT™ without applying a pressure correction. As mentioned above it was not possible to

calculate an exact pressure correction for each individual fuel due to all the tests being

conducted at the same starting pressure. A reasonable approximation can, however, be

made by applying the pressure dependence of the reaction-controlled combustion

efficiency as expressed in Equation 2.31. The shock tube results also needed to be

adjusted to account for the nitrogen dilution. This was done based on experimental and

modelling work by Zeng et al. [71] which indicated that a correction factor of 0.529 was

applicable. The shock tube ignition delays (ID843 K) were thus converted to IQT™

conditions (ID843 K’) by applying the relationship expressed in Equation 6.4 and are

compared with the IQT™-measured delays in Figure 6.4.

Chapter 6

63

𝐼𝐷843 𝐾′ = 0.529 𝐼𝐷843 𝐾 (

𝑃𝐼𝑄𝑇

𝑃𝑆𝑇)1.75

⁄ [6.4]

where 0.529 is the dilution correction factor [71]. It is interesting to note that the

constant evaporation and mixing delay offset contained in the calculation of DCN values

(Equation 6.1), was also revealed as a constant offset in the best-fit linear trend line in

the comparison with the shock tube ignition delays.

Figure 6.4: Comparison of IQT™ and shock tube ignition delays

The various ignition delay measurements were compared using a correlation matrix

analysis (see Table 6.3) which confirmed that the shock tube ignition delays at 833 K

correlated very strongly with the delays converted to IQT™ conditions (ID843 K‘).

(Absolute coefficients greater than 0.7 were considered indicative of strong correlations

while 0.5 to 0.7 were considered to be indicative of weak but significant correlations.)

The analysis also revealed a strong positive correlation between IQT™ delays and the

shock tube results at 833 K (1000/K = 1.2), but weak correlations in the high

temperature regime, 1000/K = 1.0 and 0.8. While this is not an unqualified result it

confirmed confidence in the ignition delay data sets measured in both the shock tube

and IQT™. Due to the very strong positive correlation revealed between the 1000 K and

1250 K shock tube delays which indicated that these results were essentially covariant,

only the 1000 K metric was used in the subsequent analyses. Similarly the 833 K shock

Results

64

tube delays were subsequently considered to also be a reasonable proxy representation

of results converted to the IQT™ temperature and pressure operating conditions

Table 6.3: Correlation analysis of IQT™ and shock tube ignition delay results

ID1250 K ID1000 K ID833 K ID843 K‘ IDIQT

ID1250 K (1000/T = 0.8) 1

ID1000 K (1000/T = 1.0) 0.99 1

ID833 K (1000/T = 1.2) 0.66 0.76 1

ID843 K‘ 0.67 0.78 1.00 1

IDIQT 0.40 0.51 0.86 0.85 1

-0.5 to 0.5 -0.7 to -0.5 ; 0.5 to 0.7 <-0.7 ; >0.7

6.3 Gas turbine model-combustor: LBO evaluation results

The results of LBO evaluation in the model-combustor are shown in Figures 6.5 to 6.7. It

reports the global equivalence ratios at which the blowout events were detected as a

function of the air mass flow rate for the three different combinations of burner

configuration and air pre-heat temperature.

While there were marked differences in the characteristic shape-functions associated

with the three different combustor test series, they each exhibited statistically

consistent trends for all eight test fuels. Clear differences were observed between the

eight test fuels and superficially these differences appeared to carry through

consistently over the greater part of the air mass flow range and between different test

series.

However, a transition was observed between the results at the lowest air mass flow

rates and the results at higher air mass flow rates above approximately 5 g/s. This was

broadly in line with the behaviour that was predicted by the theoretical treatment in

Section 4.3 as illustrated by Figure 4.3 which showed combustion limits being

constrained by evaporation below a critical air mass flow rate. These results will be

explored in greater detail in a subsequent discussion section.

Chapter 6

65

Figure 6.5: Model-combustor LBO results - burner configuration A, 323 K air pre-heat

Figure 6.6: Model-combustor LBO results - burner configuration A, 413 K air pre-heat

Results

66

Figure 6.7: Model-combustor LBO results - burner configuration B, 323 K air pre-heat

6.4 Gas turbine model-combustor: fuel spray evaluation results

The PDA spray measurements in the laboratory-scale model-combustor were conducted

at two stoichiometric air-fuel ratios: one at 5% rich of the LBO stoichiometry and

another at a stoichiometry of ɸ = 0.6. This allowed both investigation of the spray

behaviour as close to blowout as practical, as well as a comparison between all fuels at

exactly the same stoichiometric air-fuel ratio. For each stoichiometric set point,

measurements were performed at two air mass flow rates (2.2 g/s and 6.5 g/s) and

three distances from the nozzle exit plane (z = 15 mm, z = 25 mm , and z= 35 mm). The

combustor was traversed along the y-axis at each z-axis location in 2 mm increments.

The PDA test results are summarised in Appendix D, which reports Sauter mean

diameter (SMD), D32, values to capture the diameter distribution for each fuel,

combustion condition, and measurement position as a single representative number.

There were no significant differences in the measured behaviour between the two

stoichiometric conditions. The measurements at 6.5 g/s air mass flow rate revealed

clear fuel-specific trends that applied across all test conditions and measurement

positions. All eight test fuels displayed similar distributions over the measurement area

at the higher flow rate.

Chapter 6

67

The measurements at the lower air mass flow rate (2.2 g/s) displayed greater variability

and considerably larger droplet diameters. The spray distributions were also not

consistent across all eight test fuels, with HN deviating from the distribution displayed

by the rest of the test fuels at the lower flow rate.

Figures 6.8 and 6.9 show representative SMD results for ɸ = 0.6 and an air mass flow

rate of 6.5 g/s at the maximum and minimum axial measurement positions. Figure 6.10

reflects the SMD results for an air mass flow rate of 2.2 g/s at ɸ = 0.6 at a distance of

15 mm from the nozzle exit plane. (Note the increase in the vertical scale for Figure 6.10

relating to the lower air mass flow.)

Figure 6.8: Radial SMD profiles (ɸ = 0.6, �̇�𝑨 = 6.5 g/s, z = 15 mm from exit plane)

Results

68



69

7 Analysis and Discussion of Results

In the following discussion the results from the different aspects of the experimental

programme are combined to interrogate the influence of various fuel properties on LBO

behaviour. In contrast with the preceding sections, the spray and evaporation behaviour

of the fuels are discussed first because they are important in the evaluation of the so-

called chemical differences. The physical properties and spray behaviour were evaluated

against the theoretical combustion efficiency in an evaporation-controlled system, while

the chemical differences were examined by testing the assumption that the LBO results

had been driven by a reaction-controlled system.

7.1 Spray and evaporation-controlled system analysis of LBO results

The results of the PDA measurements were used to calculate relative SMD values at

each air mass flow condition and for both stoichiometry conditions with CJF as

references. Radial measurements were averaged for each axial measurement positions

and overall for all axial positions. The lower air mass flow rate condition did not yield

sufficient data at the z = 35 mm position for calculating relative SMD numbers. As can be

seen in Table 7.1 the average z-axis values were credible representations of the

measurements at the individual z-axis positions and there was no significant difference

in the relative SMD results between the two stoichiometry conditions. The subsequent

analyses were therefore performed using the average values for the ɸLBO+5% case as

being representative of the PDA performance at the two air mass flow rates.

The spray behaviour was evaluated by comparing the experimental PDA results with

theoretical evaporation-controlled combustion efficiencies for different fuels. Recalling

Equation 2.6, the relative evaporation-controlled combustion efficiencies for two fuels

can be expressed as Equation 7.1.



⁄ [2.6]

∴ 𝜂𝑒2𝜂𝑒1

⁄ =𝜆𝑒𝑓𝑓2

𝜆𝑒𝑓𝑓1⁄

(

1𝐷02

𝐷01⁄

)

2

[7.1]

Summary, Conclusions and Recomendations

70

Table 7.1: PDA spray measurements: Relative SMD values averaged per axial measurement position

ɸ z Air mass flow [g/s]

Test Fuel

CJF CJF/D SJF1 SJF2 SJF2+ LTSK HTSK HN

ɸLBO + 5% 15 2.2 1 0.990 1.042 1.016 0.969 0.998 0.949 1.008

ɸLBO + 5% 25 2.2 1 0.993 1.031 1.004 0.963 1.014 1.006 1.009

ɸLBO + 5% Ave 2.2 1 0.991 1.036 1.010 0.966 1.006 0.978 1.009

ɸLBO + 5% 15 6.5 1 0.987 0.910 1.005 0.971 0.924 0.880 0.929

ɸLBO + 5% 25 6.5 1 0.996 0.883 1.022 0.980 0.915 0.880 0.912

ɸLBO + 5% 35 6.5 1 0.983 0.864 1.021 0.992 0.897 0.866 0.899

ɸLBO + 5% Ave 6.5 1 0.989 0.886 1.016 0.981 0.912 0.875 0.913

0.6 15 2.2 1 0.959 1.006 1.003 0.943 0.973 0.936 0.993

0.6 25 2.2 1 1.001 1.041 1.015 0.975 1.038 1.021 1.024

0.6 Ave 2.2 1 0.980 1.024 1.009 0.959 1.005 0.978 1.008

0.6 15 6.5 1 0.996 0.915 0.997 0.959 0.910 0.885 0.920

0.6 25 6.5 1 0.998 0.885 1.017 0.986 0.911 0.866 0.889

0.6 35 6.5 1 0.995 0.869 1.023 1.000 0.893 0.856 0.903

0.6 Ave 6.5 1 0.996 0.890 1.012 0.982 0.905 0.869 0.904

The effective evaporation constants were calculated using Equation 7.2 which was

derived empirically from plots by Chin and Lefebvre that relate effective evaporation

constants to normal boiling point, droplet size, velocity, pressure and ambient

temperature [20]. They presented effective evaporation constants versus boiling point

(Tbn) for various values of droplet size-velocity products (UD0) at three pressures and

three ambient temperatures.

𝜆𝑒𝑓𝑓 ∝ 𝑃0.124 𝑇1.652

𝑇𝑏𝑛0.501 [7.2]

The 100 kPa plots were employed as representative of the atmospheric pressure

conditions under which the PDA measurements were performed. Unlike the earlier

treatment, fuel-specific volatility differences were taken into account by including the

boiling point, Tbn. The boiling point of each fuel was approximated as the distillation

temperature at which 50% is recovered. Chin and Lefebvre acknowledged that while a

single fuel property cannot fully describe evaporation characteristics of any given fuel,

average boiling point has the benefit of being directly related to vapour pressure and

fuel volatility. The PDA measurements revealed that the products of velocity and droplet

size, UD0, ranged between 150 and 1000 depending on the test conditions and

Chapter 7

71

measurement position. For any given test point, the fuel-to-fuel variation of UD0 was,

however, relatively insignificant, which implied that nominally representative values

could be used with confidence to calculate evaporation constants over the full spectrum

of test conditions. Comparison of different system temperatures revealed that while the

absolute values of the effective evaporation constants were influenced strongly, the

relative numbers were insensitive to system temperature. This enabled the use of

nominally representative values of system temperature with confidence.

The relative SMD values were modelled based on Equation 2.28. By comparing relative

SMD numbers, the influence of velocity and the characteristic dimension were

cancelled. Air/liquid ratios (ALRs) at the relevant stoichiometry conditions (ɸ = 0.6 and

ɸ = ɸLBO+5% respectively) were employed although it should be noted that the relative

results were not significantly impacted.

𝑆𝑀𝐷 ∝ (𝜎


(1 +1

𝐴𝐿𝑅) [2.28]

The results of the modelled relative effective evaporation constants, SMDs, and

evaporation-controlled combustion efficiencies are presented in Table 7.2. Due to the

use of relative normalised metrics the influence of air mass flow was largely eliminated

with only the ALR values introducing differences at the two test conditions. These

differences were, however, of the order of 1% and lower. The relative values were,

therefore, averaged over the test burner configurations, air pre-heat set points, and the

two air mass flow rates at which PDA measurement were performed.

Table 7.2: Average relative effective evaporation constants, SMDs, and evaporation-controlled combustion efficiencies of the test fuel matrix.


𝜆𝑒𝑓𝑓.(𝑟𝑒𝑙) 1 0.978 1.064 0.998 0.999 1.078 1.047 1.090

𝑆𝑀𝐷(𝑟𝑒𝑙) 1 0.996 0.970 1.008 1.003 0.953 0.955 0.967

𝜂𝑒(𝑟𝑒𝑙) 1 0.986 1.132 0.982 0.992 1.188 1.147 1.164

An analysis was performed of the linear correlation between these calculated relative

spray and evaporation metrics and the relative experimental PDA results from Table 7.1.

The results of this analysis are presented in Table 7.3 with 𝑆𝑀𝐷(𝑟𝑒𝑙) referring to the

averaged calculated relative SMD values and 𝑃𝐷𝐴(𝑟𝑒𝑙) referring to the measured

relative SMD values. The air mass flow rate conditions (2.2 g/s and 6.5 g/s) are identified


72

as “LF” (low flow) and “HF” (high flow) respectively. The two stoichiometric ratios at

which the PDA measurements were conducted are identified as “LBO” for the ɸLBO+5%

test condition and “0.6” for the ɸ = 0.6 test condition.

Table 7.3: Correlation analysis of relative spray and evaporation metrics and the relative experimental PDA results.

𝜆𝑒𝑓𝑓.(𝑟𝑒𝑙) 𝑆𝑀𝐷(𝑟𝑒𝑙) 𝜂𝑒(𝑟𝑒𝑙) 𝑃𝐷𝐴(𝑟𝑒𝑙)

(𝐿𝐹; 𝐿𝐵𝑂)

𝑃𝐷𝐴(𝑟𝑒𝑙)

(𝐿𝐹; 0.6)


(𝐻𝐹; 𝐿𝐵𝑂)


(𝐻𝐹; 0.6)

𝜆𝑒𝑓𝑓.(𝑟𝑒𝑙) 1

𝑆𝑀𝐷(𝑟𝑒𝑙) -0.874 1

𝜂𝑒(𝑟𝑒𝑙) 0.964 -0.972 1

𝑃𝐷𝐴(𝑟𝑒𝑙)(𝐿𝐹; 𝐿𝐵𝑂) 0.441 -0.198 0.320 1

𝑃𝐷𝐴(𝑟𝑒𝑙)(𝐿𝐹; 0.6) 0.520 -0.265 0.397 0.967 1

𝑃𝐷𝐴(𝑟𝑒𝑙)(𝐻𝐹; 𝐿𝐵𝑂) -0.840 0.930 -0.913 -0.231 -0.243 1

𝑃𝐷𝐴(𝑟𝑒𝑙)(𝐻𝐹; 0.6) -0.871 0.946 -0.937 -0.221 -0.255 0.996 1

-0.5 to 0.5 -0.7 to -0.5 ; 0.5 to 0.7 <-0.7 ; >0.7

The relative PDA measurements that were conducted at the higher air mass flow rate

showed a strong correlation with the calculated SMD values and strong negative

correlations with the calculated effective evaporation constants and evaporation

efficiencies. This confirmed that the measured spray behaviour was in agreement with

the theoretical treatment for a prefilming airblast atomiser. The lower air mass flow test

data revealed no correlation with the higher air mass flow data or with the theoretical

calculated values. This indicated that the spray behaviour deviated significantly from the

theoretical optimal (design) operation at the lower flow condition.

The relative PDA measurements exhibited a very high correlation between the two

stoichiometry test conditions for both air mass flow rates which enabled the use of the

data set from either one of the two test conditions as representative of the relative

spray behaviour. (The ɸLBO+5% results were used in the subsequent analyses.)

Figure 4.3 theoretically predicted the influence of evaporation efficiency on the

experimental LBO boundaries, and the experimental LBO results resembled this

behaviour to a large degree. Based on this the possibility was investigated of LBO

behaviour at the lower air mass flow rate being governed by evaporation. The operating

range of the burner, as well as the experimental noise at very low air mass flow rates,

Chapter 7

73

resulted in a very limited set of data points in the low flow region that was of interest.

This limited the value that could be gained from comparing LBO trends in this range with

theoretical evaporation efficiencies. An alternative approach was followed by

considering the postulate that the effect of evaporation efficiency resulted in a relative

offset in blowout extinction boundaries and that this would be proportional to the

differences in LBO behaviour in the lowest air mass flow rate while the higher flow

region was shown to be governed by the chemical reaction rate (Figure 4.4). The

magnitude of the “offset” was therefore represented by the numerical difference in

extinction stoichiometries at the lowest air mass flow rate of 2.2 g/s and the 8.6 g/s test

point. The results were then normalised for each test configuration by subtracting the

results of the CJF fuel in that particular configuration. The two low-preheat temperature

test configurations were chosen because they were likely to represent the more

extreme fuel evaporation situation where the fuel differences would be more readily

discerned. It was not considered to be technically defensible to include the higher pre-

heat case. These measured phi “offset” values were compared against theoretical phi

values that were calculated with Equation 7.3 which was derived from equations 2.6, 7.2

and 2.28.

𝜙 ∝ (𝑃0.124 𝑇1.652

𝑇𝑏𝑛0.501𝜂𝑒𝜎

)0.5

− 1 [7.3]

By applying an appropriate proportionality constant the correlation between theoretical

and experimental phi values could be determined. Both burner configurations revealed

positive correlations. Configuration A revealed a weaker and configuration B a stronger

positive correlation. The results from these two burner configurations were, therefore,

combined to reduce the influence of experimental scatter. The fact that a combination

of the two data sets resulted in an improved maximum correlation indicated the benefit

of reducing the influence of experimental noise as reflected by Figure 7.1 graphically.

Figure 7.2 reports the correlation at the optimal combination of the two data sets

comprising of 75% B and 25% A.

Regardless of whether the configurations are viewed individually, numerically averaged,

or the optimal combination of the data sets is used, the correlations supported the

hypothesis that the LBO behaviour at low air mass flow was governed by evaporation

efficiency.


74

Figure 7.1: Correlation of experimental and theoretical stoichiometric values (effect of proportional representation of configurations A and B)

Figure 7.2: Correlation of experimental and theoretical stoichiometric values (75% configuration B and 25% configuration A)

Chapter 7

75

7.2 Reaction-controlled system analysis of LBO results

The possibility of ignition delay differences influencing the LBO differences that were

recorded in the model-combustor was investigated by recalling the model treatment of

a well-stirred homogeneous reactor (representative of a reaction-controlled system) as

developed in Chapter 4. The relationship between the volumetric flow rate through the

combustor and the combustion efficiency (the extent to which combustion was

completed during the available residence time) are described by Equations 4.18 and

4.23 for lean and rich combustion respectively. These expressions were used to

generate Figures 4.2 and 4.3 to illustrate theoretical differences in LBO behaviour driven

by differences in AFT and proportionality constants which in turn were based on ignition

delay differences.

The model-combustor LBO results were evaluated using the same theoretical treatment

to calculate the best fit values for the proportionality constant of each test fuel. These

constants were subsequently employed in a correlation analysis with fuel properties

(including ignition delay results) and results from the preceding spray behaviour

analysis. Equation 4.18 provided the relevant expression for the relationship between

the volumetric flow rate through the combustor and the combustion efficiency under

the lean conditions that are implicit during LBO.

�̇�𝐴


(1−𝜂𝜃)𝑖𝜙𝑖−1(1−𝜂𝜃𝜙)

𝑗


The adiabatic flame temperatures at blowout would have been influenced by both the

test fuel composition and stoichiometry. For each test fuel the GCxGC based C/H ratio

was used to calculate the adiabatic flame temperature (AFT) and air fuel stoichiometry

at each LBO point. Heptane-toluene surrogates were used to match the test fuel C/H

ratios as explained previously. A regression analysis of the calculated and measured air

mass flow was then employed to determine the proportionality constant, K, for every

LBO point. The volume was based on the combustor volume, and pressure was

atmospheric, both of which were constant for all fuels. Figure 7.3 represents an example

of these proportionality constants for burner configuration B with air pre-heated to

323 K. The detailed results for all three test configurations are presented in Appendix E.


76

Figure 7.3: Proportionality constants for modelled combustor airflow based on LBO

test data - burner configuration B, 323 K air pre-heat

The calculated proportionality constants exhibited a flow dependence that would not be

expected in a homogeneous perfectly-stirred reactor. This is indicative of the

homogeneous perfectly-stirred treatment oversimplifying the conditions relevant to the

experimental spray combustor. It was not surprising and could be ascribed to the

additional evaporation, mixing and heat transfer delays that are inherent in the spray

combustor but not applicable in a stirred reactor. These various delay elements would

be influenced by air mass flow rate to different degrees. It was shown in Section 4.4 that

the mixing delay, for example, would be essentially independent of the air mass flow

rate while the evaporation delay would be inversely proportional to the square of the

air mass flow rate. As discussed previously, burner configuration A did not possess a

strong central recirculation zone as would be present in typical real aero-engine

combustors. This was in contrast to configuration B that did exhibit a well-developed

central recirculation zone which resulted in differences in the amount of air and

combustion products being recirculated into their respective primary zones at high flow

rates. Furthermore by increasing air temperature (pre-heat) air viscosity is increased

which influences the entrainment of air into the primary zone. The calculated

proportionality constants of the three burner configuration / pre-heat combinations

would therefore be expected to exhibit different degrees of dependence on air mass

Chapter 7

77

flow rate. This was explored by investigating a modified expression which incorporated

both air mass flow dependent and independent delays.

In the earlier theoretical assessment Equation 4.27 expressed combustion efficiency in

terms of the air residence time and the ignition delay:

𝜂𝜃 = (𝜌𝑉

�̇�𝐴) (𝐾𝑒−𝐸 𝑅𝑇⁄ 𝑃0.75) [4.27]

By rearranging 4.27, the reaction proportionality constant can be expressed as:

𝐾 =𝜌𝑉 �̇�𝐴⁄

𝜂𝜃𝑒−𝐸 𝑅𝑇⁄ 𝑃0.75

[7.4]

A modified proportionality constant which incorporates an additional, fixed delay term

to allow for the physical mixture preparation can then be defined by introducing a

constant delay, 𝜏𝑐, into the air residence time in a manner similar to the treatment of

the IQT™ ignition delay calibration curve (Equation 6.1) discussed in Section 6.2.

𝐾𝜃 =𝜌𝑉 �̇�𝐴⁄ − 𝜏𝑐

𝜂𝜃𝑒−𝐸 𝑅𝑇⁄ 𝑃0.75

[7.5]

∴ 𝐾𝜃

𝐾=𝜌𝑉 �̇�𝐴⁄ − 𝜏𝑐

𝜌𝑉 �̇�𝐴⁄ [7.6]

𝐾𝜃 = 𝐾(1 − 𝐶𝜃�̇�𝐴) where 𝐶𝜃 = 𝜏𝑐

𝜌𝑉 [7.7]

Modified proportionality constants were calculated for each fuel by appropriate

selection of 𝐶𝜃 to minimise the standard deviation in the higher air mass flow region

where the effects of evaporation are diminished, and one could reasonably expect the

modified proportionality constant to be independent of air mass flow rate. The results

for all three configuration / pre-heat combinations are presented in Appendix E. The

modified proportionality constants for burner configurations A and B with 323 K air pre-

heat are shown in Figures 7.4 and 7.5. Configuration A without the strong recirculation

exhibited relatively constant values at air mass flow rates above 6.5 g/s, but

configuration B retained a small measure of flow rate dependence. At the higher pre-

heat temperature, the test results with burner configuration A exhibited even greater

air mass flow dependence. The residual dependence of the modified proportionality

constants on air flow rate could possibly be ascribed to increased entrainment of

secondary air under higher flow turbulence conditions, but importantly does not

significantly impact the relative performance of the different fuels.


78

Figure 7.4: Modified proportionality constants (Kθ) for modelled combustor airflow

based on LBO test data - burner configuration A, 323 K air pre-heat

Figure 7.5: Modified proportionality constants (Kθ) for modelled combustor airflow

based on LBO test data - burner configuration B, 323 K air pre-heat

The modified and unmodified proportionality constants (as reported in Table E1,

Appendix E) were compared by conducting a correlation analysis of the averaged values

Chapter 7

79

for all three test configurations at flow rates above 8 g/s. Both modified and unmodified

proportionality constants exhibited deviations at air mass flows below 6 g/s which

agreed with the preceding conclusion that evaporation efficiency impacted the blowout

results at low air flow rates. The relative ranking of the fuel constants were similarly

achieved by normalising them to a common reference: Test Fuel CJF. The correlation

analysis results reported in Table 7.4 revealed that the normalised proportionality

constants (K) of the three test configurations correlated strongly with each other and

with their respective modified constants (Kθ). The modified constants of the two low air

pre-heat cases also correlated strongly, but the correlation between the low and high

pre-heat cases were much weaker. The K and Kθ values averaged over their three test

configurations all correlated very strongly the individual configuration rankings with the

exception of the modified constants for the higher pre-heat case of configuration A.

Based on these results the unmodified relative proportionality constants averaged over

all configurations, KAve, were selected for use as the optimal representation of the

chemical reaction rate in subsequent analyses to represent the relative LBO

performance of the fuels.

Table 7.4: Correlation analysis of normalised LBO proportionality constants (K) and modified proportionality constants (Kθ)

K (A , 323 K)

K (B , 323 K)

K (A , 413 K)

KAve

Kθ (A , 323 K)

Kθ (B , 323 K)

Kθ (A , 413 K)

KθAve

K (A , 323 K) 1

K (B , 323 K) 0.892 1

K (A , 413 K) 0.848 0.714 1

KAve 0.976 0.930 0.906 1

Kθ (A , 323 K) 0.712 0.660 0.331 0.614 1

Kθ (B , 323 K) 0.874 0.900 0.521 0.823 0.893 1

Kθ (A , 413 K) 0.873 0.742 0.990 0.922 0.380 0.571 1

KθAve 0.922 0.866 0.647 0.871 0.920 0.963 0.692 1

-0.5 to 0.5 -0.7 to -0.5 ; 0.5 to 0.7 <-0.7 ; >0.7

As a preliminary check on the self-consistency within the various experimental data, the

influence of fuel properties on LBO behaviour was interrogated by analysing the


80

correlation between the relative LBO proportionality constants (KAve), the relative spray

behaviour (SMD(rel)) and the metrics that were used to define the combustion chemistry

of the fuel matrix (LFS, IDIQT, ID1000 K, ID833 K). The results of the analysis are reported in

Table 7.5. The following conclusions were drawn from the results of this analysis:

The SMD(rel) was chosen to represent the spray behaviour at higher air mass flow

rates because of the very strong correlation with the relative PDA measurements

(as reported in Table 7.3). Under these conditions the spray was fully developed

and adhering to the theoretical treatment for a prefilming airblast atomiser.

There was absolutely no correlation between the LBO proportionality constants

and the relative spray behaviour of the different fuels which confirmed that there

was no discernable influence of evaporation or spray on the blowout results in

the high air mass flow regime.

Table 7.5: Correlation analysis of average relative LBO behaviour, fuel spray, LFS and ignition delays

KAve SMD(rel) LFS IDIQT ID833 K ID1000 K

KAve 1

SMD(rel) 0.07 1

LFS -0.08 0.74 1

IDIQT -0.71 0.22 -0.05 1

ID833 K (1000/T = 1.2) -0.87 -0.11 -0.23 0.86 1

ID1000 K (1000/T = 1.0) -0.84 -0.01 0.07 0.51 0.76 1

-0.5 to 0.5 -0.7 to -0.5 ; 0.5 to 0.7 <-0.7 ; >0.7

Similarly there was absolutely no correlation between the LBO proportionality

constants and LFS which discredited the relevance of LFS being used to predict

the LBO behaviour of different fuels. This provided clarification of the

contradictions in literature regarding the impact of LFS on LBO, that were

identified in Chapter 4.

Relative LBO behaviour, KAve, exhibited a strong inverse correlation with IQT™

ignition delays and even stronger with the two shock tube ignition delays.

However, as was also reported in Table 6.3, while IQT™ delays correlated strongly

with the lower intermediate temperature shock tube ignition delays, the

correlation was weaker with the higher temperature delays. Viewed in

Chapter 7

81

combination these results would suggest that while the IQT™ is providing a

reasonable indication of the LBO performance, it could possibly provide an even

better metric if the IQT™ apparatus were operated at a higher temperature.

Figure 7.6: Correlation of relative LBO proportionality constants (KAve) and

intermediate temperature shock tube ignition delay (ID833 K)

7.3 Interpretation of the experimental results

The experimental programme was designed to interrogate the hypothesis that jet fuel

properties relating to evaporation and reaction behaviour could have a detectable and

significant influence on high altitude extinction behaviour. Analysis of the experimental

results provided the necessary data required to perform this task and were interpreted

by recalling the theoretical operating envelope, demarcated by the reaction and

evaporation-controlled extinction limits that was presented in Figure 2.6. The test

programme was not designed to yield the absolute values of the reaction and

evaporation-controlled extinction limits; however, by examining the relative behaviour

of different fuels it was possible to explain the altitude extinction results that were

observed during the FSJF certification tests.

First the relative LBO behaviour in the model-combustor was used to model the

theoretical relative reaction-controlled operating envelope. The proportionality

constant of the crude-derived jet fuel CJF, K’CJF, was chosen such that the extinction


82

boundary (based on Equation 2.31) passed through the blowout extinction point that

was recorded by the crude-derived test fuel (AVTUR) during the Rolls-Royce test

campaign: Mach 0.99 at 10760 m (35300 ft.) altitude [13]. The fuel-specific relative LBO

proportionality constants (KAve) that were calculated in the preceding section were

employed in Equation 7.8 to produce the relative extinction limits for the rest of the test

fuel matrix.

𝜂𝜃 ∝ 𝑉𝑒−𝐸 𝑅𝑇⁄ 𝜌1.75 �̇�𝐴⁄ [2.31]

𝜂𝜃 = 𝐾′𝐶𝐽𝐹 𝐾𝐴𝑣𝑒 𝑉 𝑒−𝐸 𝑅𝑇⁄ 𝜌1.75 �̇�𝐴⁄ [7.8]

The results are shown graphically in Figure 7.7 resulting in a number of observations

that can be made. As mentioned, test fuel SJF1 was nominally identical to the synthetic

test fuel that was employed in the FSJF certification tests (SFSAK). A blowout extinction

limit of approximately Mach 0.95 at 10270 m (33700 ft.) altitude was reported for the

SFSAK fuel [13]. The modelled extinction behaviour of SJF1 confirmed this impaired

extinction limit relative to the crude-derived reference CJF/ AVTUR with its extinction

boundary passing through Mach 0.95 at an altitude of 10150 m. (The AVTUR and SFSAK

blowout extinction points are included in figures 7.7 to 7.9 to contextualise the results.)

It is interesting to note that the relative proportionality constants for SJF1 represented a

9 to 12 % reduction in air mass flow at blowout which compares well with the

observation from the altitude test report which stated an almost 6% difference in

relative air flow at first sector blowout and more than 10% difference in final sector

blowout.

The modelled performance of the production synthetic jet fuel formulation, SJF2, was

marginally better than the reference CJF while the low-temperature FT kerosene, LTSK

achieved the highest extinction boundaries.

Chapter 7

83

Figure 7.7: Theoretical relative reaction-controlled extinction limits

Similarly a theoretical relative evaporation-controlled operating envelope was modelled

based on relative spray behaviour and relative effective evaporation constants, again

using CJF as the reference. Incorporating SMD(rel) and the effective evaporation

constants from Equation 7.2 into Equation 2.6 yields Equation 7.9 which was used to

model the relative evaporation-controlled extinction limits. The proportionality constant

K”CJF was again chosen such that the extinction boundary of CJF passed through the

blowout extinction point that was recorded by the AVTUR fuel during the Rolls-Royce

test campaign: Mach 0.99 at 10760 m (35300 ft.) altitude. The resultant extinction

boundaries for all eight test fuels are reported in Figure 7.8.

𝜂𝑒 = 𝐾𝐶𝐽𝐹" 𝑃1.124 𝑇0.652

𝑇𝑏𝑛0.501 𝑆𝑀𝐷𝑟𝑒𝑙

2 �̇�𝐴⁄ [7.9]


84

Figure 7.8: Theoretical relative evaporation-controlled extinction limits

The use of T50 as representative of the boiling point resulted in potentially exaggerated

spread in the relative performance of the different fuels, but it does not undermine the

modelled trends and relative performance.

The modelled extinction boundary of test fuel SJF1 was considerably improved over the

reference CJF. The production synthetic jet fuel formulation, SJF2, was marginally worse

than the reference CJF while the low-temperature FT kerosene, LTSK achieved the

highest extinction boundaries. The SFSAK blowout extinction point was not supported

by the evaporation-controlled boundary of SJF1. This clearly indicates that the location

of the impaired SFSAK extinction point was not associated with an evaporation-

controlled limit.

The relative spray behaviour and evaporation-controlled boundary of SJF1 also agreed

with the observation made by the Honeywell test report that the atomisation of the

FSJF was as good as or better than the Jet A reference fuel [15].

Interestingly SJF2 recorded impaired extinction limits relative to SJF1 in spite of

conforming to the distillation gradient restriction that was introduced partially in

reaction to the FSJF (SJF1) altitude relight results.

Chapter 7

85

The flash point modified SJF2+ reported improved evaporation limited extinction

boundaries relative to SJF2, but it was primarily ascribed to the relative SMD behaviour

of the two fuels with the distillation only playing a secondary role. The apparent

sensitivity to distillation could be increased slightly by using the flash point instead of T50

as the representative boiling point temperature in Equation 7.9. This would widen the

difference in relative performance of the two fuels slightly (as would be expected) and

would result in SJF2+ recording marginally better extinction limits than CJF, but it was

not considered to be a technically justified representation.

It can therefore be concluded that the reaction-controlled extinction limits were in

agreement with the altitude relight differences that were observed during the Rolls-

Royce FSJF tests, while the evaporation-controlled extinction limits were in direct

opposition to the FSJF altitude extinction results. The relative atomisation results of SJF1

were in fact in agreement with the atomisation results during the FSJF certification

tests: The atomisation of the FSJF was reported to be as good as or better than the Jet A

reference fuel for both airblast and pressure type fuel atomisers.

These results indicate that the relative impact of the reaction and evaporation efficiency

limits on altitude extinction blowout were such that the reaction-controlled limits

dominated and limited the overall altitude-Mach number envelope. Figure 7.9 compares

both evaporation- and reaction-controlled extinction limits for four of the test fuels. As

before, the proportionality constants were chosen such that the extinction boundaries

of CJF passed through the blowout extinction point of the AVTUR test fuel. The relative

performance of the other test fuels then serves to illustrate the relative impact of their

evaporation and reaction boundaries.

It is interesting to examine the effect of introducing SJF2 to compensate for the lower

blowout extinction performance recorded by SJF1. In spite of the focus on distillation

gradient differences, a chemical ignition delay improvement resulted in the reaction-

controlled extinction behaviour being improved to marginally better than the reference

CJF. The evaporation-controlled extinction limits were, however, impaired and suggests

that overall blowout extinction limits could be compromised at higher flight Mach

numbers in excess of Mach 0.85 where the evaporation boundary drops below the

reaction boundary. Consistent with the argument that was offered in the original

certification report, this effect appears beyond the typical flight envelope and the

resultant blowout extinction limits are projected to lie between the AVTUR and SFSAK


86

results. It should also be noted that it is unlikely that both evaporation and extinction

boundaries would have come into play simultaneously with the AVTUR which means

that the absolute evaporation limits for all the test fuels could potentially be greater.

The largely linear paraffin structure of LTSK has been associated in literature with

shorter ignition delays and higher DCN values. Compared to the rest of the test fuel

matrix, LTSK recorded the best performance in both evaporation-limited and reaction-

limited efficiency boundaries by significant margins. This is of particular interest due to

LTSK nominally resembling a fuel that could also be produced via a renewable synthetic

jet fuel manufacturing process.

Figure 7.9: Theoretical relative evaporation- and reaction-controlled extinction limits

87

8 Summary, Conclusions, and Recommendations

This study investigated the influence of alternative jet fuel properties on aviation gas

turbine performance at threshold combustor operating conditions. It was motivated by

previously unexplored anomalous altitude extinction results that were observed during

an industry evaluation of synthetic jet fuel as well as inconsistencies in literature dealing

with this subject.

The research methodology was based on the use of a fundamental theoretical

treatment to interpret results generated by an experimental design which adopted a

novel approach of formulating a test fuel matrix to reflect real-world alternative fuel

compositions while still enabling a targeted evaluation of the influences of both physical

and chemical reaction properties. A comprehensive characterisation was performed of

all the test fuels detailing their physical properties and chemical reaction properties. The

extinction and spray behaviours of the fuels were then evaluated in a laboratory scale

combustor and the various experimental data sets were interpreted within the context

of the theoretical analyses.

The relative performance of alternative jet fuel formulations under laboratory burner

conditions was translated to predict relative real world altitude performance. This

approach was validated against results from the abovementioned industry evaluation

and demonstrated to be very credible. This study enabled the following conclusions to

be drawn:

All three project hypotheses were substantiated.

Hypothesis 1: While most of the eight test fuel blends conformed to the legislated

physical jet fuel specifications, the non-conforming blends were deliberately

included for technical reasons. The spectrum of blends successfully achieved

significant independent variation in properties relating to evaporation and chemical

reaction behaviour.

Hypothesis 2: The test fuels recorded significant differences in spray formation,

evaporation, and blowout extinction which resulted in significant and measurable

combustion efficiency and blowout differences. Through the application of an

appropriate theoretical model these differences were shown to impact altitude

blowout results which were validated by comparison with real altitude test results.

Summary, Conclusions, and Recommendations

88

Hypothesis 3: Chemical ignition delays measured in a shock tube were shown to

correlate strongly with altitude blowout performance. The chemical ignition delays

measured in an IQT™ recorded slightly weaker correlations with LBO performance,

but also provided an accurate metric of the relative blowout extinction behaviour of

different fuels.

A critical analysis of the test data provided a technically defensible explanation

for the altitude relight results that were observed in the Rolls-Royce test

campaign that was conducted during the FSJF certification process. A credible

case has been made for the relative extinction limits being ascribed to the

relative chemical ignition delays of the fuels.

The fuel spray evaluation discredited the probability of volatility (distillation

profile) and fuel physical properties playing a significant role in the impaired

altitude performance that was recorded during the Rolls-Royce tests.

DCN ignition delays were shown to correlate strongly with the LBO results obtained in a technically relevant gas turbine model-combustor supporting its use as a chemical reaction rate proxy.

The 1000 K shock tube ignition delays recorded even stronger correlations with the LBO results.

Laminar flame speed was shown not to correlate with LBO and no support was found for its use as a proxy for reaction rate.

Evaporation-controlled combustion efficiency was shown to become a significant

factor at low air mass flow rates or when the fuel evaporation is degraded.

Different synthetic jet fuel formulations recorded wide ranging results with low

temperature Fischer Tropsch GTL kerosene (LTSK) recording the best

performance in both evaporation-limited and reaction-limited efficiency

boundaries by significant margins.

The results of this study support the importance of chemical ignition delays being

considered when evaluating the blowout behaviour of alternative jet fuel

formulation options.

Based on these conclusions, the following recommendations are made:

It is recommended that DCN values (based on the IQT™-measured ignition

delays) provide a practical and convenient metric for provisionally ranking jet fuel

ignition delay and for highlighting potential altitude extinction concerns.

Chapter 8

89

It is recommended that the development of a higher temperature version of the

IQT™ be investigated to provide an even more representative screening tool. This

could potentially be implemented more readily than shock tube testing which

would be prohibitively complex for wider industry use.

Future certification processes should afford due consideration to the chemical

ignition delay character of a candidate fuel.

A.1

Appendix A: Evaporation-controlled System

In a system where the mixing and reaction rates are sufficiently fast for the fuel

evaporation to be the rate-controlling step, the fuel is assumed to instantly mix and

burn with the surrounding air as soon as it evaporates. Under such conditions the

evaluation of evaporation and droplet lifetime is of interest since it determines the

residence time required for combustion to occur.

The evaporation of a spherical droplet is accepted to consist of an unsteady state or

heat-up period during which the droplet temperature increases with time until the drop

attains its wet-bulb temperature. This transient period is followed by relatively steady-

state evaporation during which the droplet diameter, D, changes according to the so-

called “D2 law of droplet evaporation” that was first formulated by Godsave [17]. Typical

assumptions include: that the droplet is spherical, that the fuel is a pure single boiling

point liquid and that radiant heat transfer is negligible. The mathematical expression of

the law is provided in equation A.1 where λ represents the evaporation constant, t the

time elapsed and D and D0 the instantaneous and starting droplet diameters

respectively [18].

𝐷𝑜2 − 𝐷2 = 𝜆𝑡 [A.1]

Figure A.1 shows this relationship between D2 and time as generated by Wood et al. for

kerosene and JP-4 droplets [19]. The evaporation constant corresponds to the slope of

the graph which is very small during the initial heat-up stage while the majority of the

heat supplied is used to raise the droplet temperature. As the liquid temperature rises

the slope gradually increases with time until a stage is reached where all the heat

transferred to the droplet is used as heat of vaporization and the droplet attains a

steady-state at the wet-bulb temperature. Beyond this the slope of the curve remains

virtually constant for the remainder of the droplet’s life. Figure A.1 is based on

measurements that were conducted at atmospheric pressure and 2000 K. under which

conditions the initial heat-up period only constitutes a small portion of the total

evaporation time. This is not necessarily the case for many fuels under high ambient

temperature and pressure conditions where the unsteady heat-up period can be

significant [18, 72].

Evaporation-controlled System

A.2

Figure A.1: Evaporation rate curves for

kerosene and JP-4 [19] Figure A.2: Drop size vs time ignoring heat-up

From equation A.1 it follows that the evaporation constant, or slope of the curves in

Figure A.1, are expressed by

𝜆 = 𝑑

𝑑𝑡 (𝐷2) [A.2]

𝜆 = 2𝐷𝑑

𝑑𝑡 (𝐷) [A.3]

Figure A.3: Droplet shell

The evaporation of a fuel droplet and consumption of fuel during the steady burning of

a stationary droplet in an oxidizing atmosphere is assumed to occur at a spherical

surface that is established around the droplet. The volume, V, of this droplet shell is

given by:

Appendix A

A.3

𝑉 = 𝑆𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 . 𝑑 (𝐷

2) = 𝜋𝐷2 𝑑(

𝐷

2) [A.4]

The rate of volume and mass change due to evaporation can therefore be expressed as

𝑑𝑉

𝑑𝑡= �̇� =

𝜋

2𝐷2

𝑑

𝑑𝑡(𝐷) [A.5]

𝑑𝑚

𝑑𝑡= �̇� =

𝜋

2𝜌𝐹𝐷

2 𝑑

𝑑𝑡(𝐷) [A.6]

Where 𝜌𝐹 is the fuel density. By substituting the evaporation constant from A.3 the

volume and mass flow rates simplifies to

�̇� = 𝜋

2𝐷2

𝜆

2𝐷=

𝜋

4𝐷𝜆 [A.7]

�̇� = 𝜋

4𝜌𝐹𝐷𝜆 [A.8]

The mass change between the initial droplet diameter (D = D0 ; m = m0) and droplet

extinction (D = 0 ; m = 0) can be derived by integration of A.6.

𝑑𝑚 = 𝜋

2𝜌𝐹𝐷

2 𝑑(𝐷) [A.9]

∫ 𝑑𝑚0

𝑚0= ∫

𝜋

2𝜌𝐹𝐷

2 𝑑(𝐷)0

𝐷0 [A.10]

𝑚0 = 𝜋

2𝜌𝐹 [

𝐷𝑜

3

3 ] =

𝜋

6𝜌𝐹𝐷0

3 [A.11]

By assuming λ to be constant, integration of A.2 yields the droplet lifetime, te.

𝑡𝑒 = 𝐷02

𝜆⁄ [A.12]

The evaporation constant was defined by Godsave for steady-state evaporation under

quiescent conditions. Chin and Lefebvre defined an effective evaporation constant, λeff,

to include the effects of the heat-up period and forced convection [20].

𝑡𝑒 = 𝐷02

𝜆𝑒𝑓𝑓⁄ [A.13]

The average rate of droplet evaporation, �̇�𝑎𝑣𝑒, follows from dividing the droplet mass

evaporated by the droplet lifetime.

𝑚0

𝑡𝑒=

𝜋

6𝜌𝐹 [

𝐷𝑜

𝑡𝑒

3 ] =

𝜋

6𝜌𝐹 𝐷0 [

𝐷𝑜

𝑡𝑒

2 ] [A.14]

∴ �̇�𝑎𝑣𝑒 = 𝜋

6𝜌𝐹 𝐷0𝜆𝑒𝑓𝑓 [A.15]


A.4

The average evaporation rate of n fuel drops with mean initial diameter of D0 contained

in a volume of air, can be expressed as

�̇�𝑎𝑣𝑒 = 𝜋

6𝑛𝜌𝐹 𝐷0𝜆𝑒𝑓𝑓 [A.16]

By assuming that the mass of fuel is equivalent to the droplet mass multiplied by the

number of droplets (n x 2.12), the fuel/air ratio, q, in the volume can be expressed as

𝑞 = 𝑚𝐹

𝑚𝐴⁄ = 𝑛𝜋

6𝜌𝐹𝐷0

3

𝑉 𝜌𝐴⁄ [A.17]

The number of droplets in the volume can therefore be expressed in terms of the

fuel/air ratio, densities and diameter (A.18) and substituted into A.16 to yield the

average rate of evaporation of a fuel spray.

𝑛 = 6𝑞𝑉𝜌𝐴

𝜋𝜌𝐹𝐷03⁄ [A.18]

�̇�𝑎𝑣𝑒 = 𝜋

6𝜌𝐹 𝐷0𝜆𝑒𝑓𝑓 .

6𝑞𝑉𝜌𝐴𝜋𝜌𝐹𝐷0

3⁄ = 𝜌𝐴𝜆𝑒𝑓𝑓𝑉𝑞

𝐷02⁄ [A.19]

This expression for the average fuel spray evaporation rate can be employed in

determining the combustion efficiency of a system in which the mixing and reaction

rates are sufficiently fast for the fuel evaporation to be the rate-controlling step. The

fuel is assumed to instantly mix and burn with the surrounding air as it evaporates. The

combustion efficiency is therefore governed by the ratio of the rate of fuel evaporation

within the combustion zone to the rate of fuel supply.

𝜂𝑒 = �̇�𝑎𝑣𝑒

𝑞𝑜𝑣�̇�𝐴⁄ =

�̇�𝑎𝑣𝑒𝑓𝑐𝑞𝑐�̇�𝐴⁄ [A.20]

Where fc is the fraction of the total combustor airflow that takes part in combustion and

qov and qc are the overall and combustion zone fuel/air ratios respectively. Substituting

A.19 into A.20 yields

𝜂𝑒 = 𝜌𝐴𝜆𝑒𝑓𝑓𝑉𝑞

𝐷02𝑓𝑐𝑞𝑐�̇�𝐴

⁄ = 𝜌𝐴𝜆𝑒𝑓𝑓𝑉


⁄ [A.21]

Appendix A

A.5

From [A.13] the effective evaporation constant can be expressed in terms of the

evaporation time, te: 𝜆𝑒𝑓𝑓 = 𝐷02

𝑡𝑒⁄ .



⁄ = 𝜌𝐴𝑉

𝑡𝑒𝑓𝑐�̇�𝐴⁄ [A.22]

Since 𝜌𝐴𝑉 is the mass of air in the combustor, 𝜌𝐴𝑉

𝑓𝑐�̇�𝐴⁄ = ∆𝑡𝑟 = the air residence time

in combustor, which means that A.22 agrees with the logical argument that the

combustion efficiency amounts to the ratio of residence time of the air in the combustor

to the evaporation time for the droplets.

𝜂𝑒 = ∆𝑡𝑟

∆𝑡𝑒⁄ [A.23]

Equation A.21 expresses combustion efficiency in terms of the combustor operating

conditions (𝜌𝐴 , 𝜆𝑒𝑓𝑓 , �̇�𝐴), physical combustor dimensions (V), fuel spray characteristics

(D0) and fuel properties (D0 , 𝜆𝑒𝑓𝑓). The equation, however, allows the evaporation

efficiency to assume a value greater than 100%. This occurs when the time required for

complete evaporation is less than the available time and the fuel within the primary

recirculation zone is thus fully vaporised. Under these conditions the combustion

efficiency should be assigned a value of 100%. Lefebvre proposes the use of a modified

form similar to A.24 in order to avoid the abrupt discontinuity [21]. The two expressions

are compared on the basis of combustion air residence in Figure A.4.

𝜂𝑒 = 1 − exp (−𝐵𝜌𝐴𝜆𝑒𝑓𝑓𝑉


⁄ ) [A.24]


A.6

Figure A.4: Comparison of the direct (eq. A.21) and exponential (eq. A.24) expressions

for evaporation efficiency

B.1

Appendix B: Comprehensive Fuel Property Analysis Results

Property Units Limits CJF CJF/D SJF1 SJF2 SJF2+ LTSK HTSK HN

Appearance Appearance Visual Report PASS PASS PASS PASS PASS PASS PASS PASS Colour, Saybolt - Report 18 24 >+30 >+30 >+30 >+30 >+30 >+30 Particulate Contaminants

mg/l 1.0 max 0.6 0.3 2.1 0.2 <0.1 0.1 0.3 0.5

Particulate >= 4 µm (c) Report 1423.0 1571.9 19605.9 194.3 1105.0 193.3 120.8 472.4 >= 6 µm (c) Report 496.3 409.2 7636.2 82.3 422.6 62.9 69.9 16.0 >= 14 µm (c) Report 79.9 30.2 375.6 10.5 61.7 6.5 18.2 4.0 >= 21 µm (c) Report 29.1 8.5 36.2 2.4 16.5 1.9 6 1.6 >= 25 µm (c) Report 15.1 3.6 10.3 1 7.4 1.1 2.8 0.9 >= 30 µm (c) Report 6.5 1.3 3.1 0.4 3.2 0.2 1.1 0.4 Composition Total Acidity mgKOH/g 0.015 max 0.002 0.001 0.002 0.001 <0.001 0.003 0.001 0.001 Total Aromatics vol % 26.5 max 17.1 8.6 13.0 10.2 9.5 0.0 0.8 6.1 Paraffins vol % - 82.0 90.3 84.2 86.7 86.4 99.5 97.7 93.2 Olefins vol % - 0.9 1.1 2.8 3.1 4.1 0.5 1.6 0.8 Total Sulphur mass% 0.30 max 0.12 0.07 <0.01 <0.01 <0.01 <0.01 <0.01 <0.01 Sulphur Mercaptan mass% 0.0030 max 0.0016 0.0009 <0.0010 0.0010 <0.0010 <0.0010 <0.0010 <0.0010 Antioxidant mg/l - <4 <4 <4 16 <4 4 <4 <4 API Gravity degAPI - 46.0 50.6 45.6 41.8 42.4 61.0 54.2 54.6 Combustion Specific Energy MJ/kg 42.80 min 43.32 43.70 43.30 43.29 43.30 44.30 43.97 43.80 Smoke Point mm 25.0 min 28.0 32.0 29.0 25.0 25.0 28.0 28.0 30.0 DCN ex IQT™ Report 45.9 58.5 32.4 30.9 30.6 62.0 30.6 52.6 JFTOT Control temperature °C 260 min - - - - - - - - Filter Pressure Diff mmHg 25.0 max 0 0 0 0 0 0 0 0 Tube Deposit Rating <3 1 <1 <1 <1 <1 <1 <1 <2 Filter Test - F F F F F F F F

Comprehensive Fuel Property Analysis Results

B.2


Volatility Initial Boiling Point °C Report 150.4 162.6 161.6 168.1 160.7 147.3 166.0 160.8 5% °C 164.3 181.2 163.4 173.4 169.3 154.8 169.9 163.6 10% °C 205.0 max 170.4 187.6 163.1 176.9 171.0 156.1 171.2 163.7 20% °C 176.9 193.6 165.9 180.5 178.6 160.1 172.5 164.0 30% °C 182.3 198.3 168.0 185.2 184.1 162.6 173.9 164.5 40% °C 188.6 202.0 170.2 190.5 189.7 165.2 175.7 165.1 50% °C Report 195.7 205.6 172.7 196.6 196.2 168.8 178.0 165.7 60% °C 203.5 209.0 176.5 204.3 204.1 172.9 180.8 166.6 70% °C 212.2 211.7 181.0 214.2 214.2 177.4 184.5 167.7 80% °C 222.1 215.0 186.6 226.9 227.5 182.3 189.3 169.2 90% °C Report 235.1 221.6 195.7 241.5 242.2 188.0 196.4 172.3 95% °C 245.0 232.2 204.7 249.3 250.8 191.7 202.5 178.0 Final Boiling Point °C 300.0 max 259.9 250.9 225.2 259.3 258.5 195.2 215.0 204.9 Recovery vol % 97.6 98.0 95.8 98.1 98.3 98.1 98.1 98.0 Residue vol % 1.5 max 1.0 1.0 1 1.0 1.0 0.9 0.9 0.8 Loss vol % 1.5 max 1.0 1.0 0.8 0.9 0.7 0.9 0.9 1.2 T50-T10 °C > 15 25.3 18.0 9.6 19.7 25.2 12.7 6.8 2.0 T90-T10 °C > 40 64.7 34.0 32.6 64.6 71.2 31.9 25.2 8.6 Flash pt. @ 101,3 kPa °C 38.0 - 50.0 42.5 47.0 42.5 53.5 38.0 38 46.5 47.0 Density at 15 °C g/mL 0.7750 -

0.8400 0.7961 0.7755 0.7979 0.8152 0.8128 0.7334 0.7604 0.7587

Density at 20 °C g/mL 0.7710 - 0.8360

0.7931 0.7725 0.7949 0.8122 0.8098 0.7304 0.7574 0.7557

Fluidity Freezing Point °C -47.0 max -50.4 -19.9 -71.2 -57.0 -57.3 -50.7 <-56.8 -45.3 Viscosity @ -20 °C mm2/s 8.0 max 3.99 4.30 3.12 5.23 5.51 2.58 3.68 2.48 Viscosity @ 40 °C cSt - 1.20 1.29 1.05 1.46 1.42 0.91 1.16 0.91

Appendix B

B.3


Contaminants Existent gum mg/100

ml 7 max 2.4 0.6 0.8 0.9 0.1 1.2 1.0 0.8

Water content mg/kg 80 max 43 37 21 33 20 39 23 12 Free water - <30 <30 <30 <30 <30 <30 <30 <30 Fuel with Static Dissipator Additive

70 min 76 84 - - - - - -

Fuel without Static Dissipator Additive

85 min - - 89 98 99 98 94 98

Conductivity Electrical Conductivity pS/m 50 - 450 97 40 33 1 1 1 14 1 Lubricity BOCLE, WSD mm 0.85 max 0.68 0.71 0.80 0.71 0.72 0.69 0.70 0.83 Corrosion Copper Corrosion 1 max 1A 1A 1A 1A 1B 1A 1A 1B Surface Tension S. Tension @ 25°C mN/m 25.68 25.61 24.06 25.97 25.94 23.45 23.85 24.25 S. Tension @ 40°C mN/m 24.3 24.1 22.8 24.69 24.47 22.07 22.16 22.78

C.1

Appendix C: Shock Tube Ignition Delay Results

Figure C.1: Test fuel CJF: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Figure C.2: Test fuel CJF/D: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Shock Tube Ignition Delay Results

C.2

Figure C.3: Test fuel SJF1: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Figure C.4: Test fuel SJF2: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Appendix C

C.3

Figure C.5: Test fuel SJF2+: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Figure C.6: Test fuel LTSK: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Shock Tube Ignition Delay Results

C.4

Figure C.7: Test fuel HTSK: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Figure C.8: Test fuel HN: Shock tube measured (markers) and modelled (curve) ignition delay results based on eq. 6.3.

Appendix C

C.5

Table C.1: Arrhenius function coefficients for modelled ignition delay

Fuel Ln(A1) B1 Ln(A2) B2 Ln(A3) B3

CJF -17.262 8714 2.22 -4808.1 -12.15 11051 CJF/D -17.262 8714 1.42 -4808.1 -12.07 11051 SJF1 -17.262 8714 3.68 -4808.1 -11.98 11051 SJF2 -17.262 8714 3.25 -4808.1 -12.24 11051 SJF2+ -17.262 8714 2.78 -4808.1 -12.25 11051 LTSK -17.262 8714 1.92 -4808.1 -12.41 11051 HTSK -17.262 8714 3.43 -4808.1 -12.02 11051 HN -17.262 8714 1.81 -4808.1 -12.34 11051

D.1

Appendix D: PDA Spray Measurement Results

A selection of results from the PDA spray measurement that were conducted in the

laboratory-scale model-combustor are summarised here as figures D.1 to D.12. Each

figure represents a set combustion condition at a set stoichiometry, air mass flow rate

and axial distance from the exit plane (z-axis). Measurements for each combustion

condition were performed at radial (y-axis) increments of 2 mm. All tests were

conducted under stable combustion conditions and at an air pre-heat temperature of

323 K. Two stoichiometric air-fuel ratios were evaluated: one at the LBO stoichiometry

plus 5% and another at a stoichiometry of ɸ = 0.6. This allowed both investigation of the

spray behaviour as close to blowout as practical, as well as a comparison between all

fuels at exactly the same stoichiometric air-fuel ratio. Sauter mean diameter, D32, (SMD)

values are reported to capture the diameter distribution for each fuel, combustion

condition, and measurement position as a single representative number.

PDA Spray Measurement Results

D.2

Table D.1 reports the relative SMD values for each air mass flow condition and for both

stoichiometric conditions. Radial measurements were averaged for each axial

measurement positions and overall for all axial positions. The lower air mass flow rate

condition did not yield sufficient data at the z = 35 mm position for calculating relative

SMD numbers.

Table D.1: PDA spray measurements: Relative SMD values averaged per axial measurement position

ɸ z Air mass

flow [g/s]

Test Fuel


ɸLBO + 5% 15 2.2 1 0.990 1.042 1.016 0.969 0.998 0.949 1.008

ɸLBO + 5% 25 2.2 1 0.993 1.031 1.004 0.963 1.014 1.006 1.009

ɸLBO + 5% Ave 2.2 1 0.991 1.036 1.010 0.966 1.006 0.978 1.009

ɸLBO + 5% 15 6.5 1 0.987 0.910 1.005 0.971 0.924 0.880 0.929

ɸLBO + 5% 25 6.5 1 0.996 0.883 1.022 0.980 0.915 0.880 0.912

ɸLBO + 5% 35 6.5 1 0.983 0.864 1.021 0.992 0.897 0.866 0.899

ɸLBO + 5% Ave 6.5 1 0.989 0.886 1.016 0.981 0.912 0.875 0.913

0.6 15 2.2 1 0.959 1.006 1.003 0.943 0.973 0.936 0.993

0.6 25 2.2 1 1.001 1.041 1.015 0.975 1.038 1.021 1.024

0.6 Ave 2.2 1 0.980 1.024 1.009 0.959 1.005 0.978 1.008

0.6 15 6.5 1 0.996 0.915 0.997 0.959 0.910 0.885 0.920

0.6 25 6.5 1 0.998 0.885 1.017 0.986 0.911 0.866 0.889

0.6 35 6.5 1 0.995 0.869 1.023 1.000 0.893 0.856 0.903

0.6 Ave 6.5 1 0.996 0.890 1.012 0.982 0.905 0.869 0.904

Appendix D

D.3

Figure D.1: Radial SMD profiles (ɸ = ɸLBO+5%, �̇�𝑨 = 2.2 g/s, z = 15 mm from exit plane)



D.4



Appendix D

D.5




D.6

Figure D.7: Radial SMD profiles (ɸ = 0.6, �̇�𝑨 = 2.2 g/s, z = 15 mm from exit plane)


Appendix D

D.7




D.8



E.1

Appendix E: Modelled LBO Air Flow Proportionality Constants

Table E1: Normalised LBO proportionality constants and modified LBO proportionality constants


K (A , 323 K) 1.00 1.04 0.88 1.00 0.99 1.08 0.91 0.97

K (B , 323 K) 1.00 1.12 0.93 1.04 1.02 1.09 0.95 1.05

K (A , 413 K) 1.00 0.99 0.93 1.02 0.98 1.12 0.95 1.02

KAve 1.00 1.05 0.91 1.02 1.00 1.10 0.94 1.01

Kθ (A , 323 K) 1.00 1.06 0.81 1.04 0.94 0.93 0.83 0.85

Kθ (B , 323 K) 1.00 1.09 0.88 1.01 0.99 1.02 0.89 0.96

Kθ (A , 413 K) 1.00 0.99 0.93 1.02 1.00 1.12 0.93 1.02

KθAve 1.00 1.05 0.87 1.02 0.98 1.02 0.88 0.94

Modelled LBO Air Flow Proportionality Constants

E.2

Figure E1: Proportionality constants for modelled combustor airflow based on LBO

test data - burner configuration A, 323 K air pre-heat


test data - burner configuration A, 413 K air pre-heat

Appendix E

E.3


test data - burner configuration B, 323 K air pre-heat

Figure E4: Modified proportionality constants (Kθ) for modelled combustor airflow


Modelled LBO Air Flow Proportionality Constants

E.4




based on LBO test data - burner configuration B, 323 K air pre-heat

References

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[8] D. Fyffe, J. Moran, K. Kannaiyan, R. Sadr, and A. Al-Sharshani, “Effect of GTL-like jet fuel composition on GT engine altitude ignition performance – Part I: Combustor operability,” GT2011-45487, in Proceedings of ASME Turbo Expo 2011, Vancouver, Canada, 2011.

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References

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[49] G. Damkoehler, Z. Elektrochem, 46, 601, 1940.

[50] K.J. Laidler, The development of the Arrhenius equation, in Journal of Chemical Education, 61 (6), 1984, pp.494-498.

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Combustion in Aviation Gas Turbine Engines of University

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